NACA airfoil blade reverse modeling method, device, computer equipment and storage medium

Abstract

The present disclosure provides a method, apparatus, computer device, and storage medium for reverse modeling of a NACA airfoil blade. The method comprises scanning a NACA airfoil blade to be identified to obtain a blade solid model; dividing the blade solid model into regions, and calculating corresponding airfoil shaping parameters based on the area values ​​of each region; outputting the model of the NACA airfoil blade to be identified based on the airfoil shaping parameters; and performing a lofting operation on the cross-sectional profile based on the model to generate a NACA airfoil blade model. By using the area method to divide the blade solid model into regions, the corresponding airfoil shaping parameters can be calculated based on the regional area distribution. Finally, the model can be determined based on the airfoil shaping parameters, reducing the error caused by the blade profile. Only the blade area needs to be calculated, which reduces the amount of data processing, effectively improving the accuracy of model identification, and reducing the data processing difficulty of reverse modeling of airfoil blades.

Description

Technical Field

[0001] The present disclosure relates to the technical field of airfoil blades, and in particular to a method, apparatus, computer equipment, and storage medium for reverse modeling of NACA airfoil blades. Background Art

[0002] Blades are key components of wind turbines and hydro turbines. NACA airfoil blades, due to their unique structure, are widely used in both types of machinery. However, with their widespread use, blade damage caused by cavitation and erosion has increased. This issue requires reverse modeling and redesign of these blades. Furthermore, reverse modeling of NACA airfoil blades can help reduce blade maintenance costs and improve design and manufacturing efficiency.

[0003] The current research on reverse design of blades is as follows:

[0004] 1. CN117669229A - A blade reverse modeling method and system based on airfoil parameter extraction. On the one hand, a blade reverse modeling method is proposed. The blade is first scanned to obtain a blade polyhedron model, and then the missing parts of the obtained model are filled. After filling, the filled area is smoothed to obtain a smoothed model. The smoothed model is aligned with the coordinate system by taking the top plane of the smoothed model as the "xoz" plane, the tangent point of the leading edge of the airfoil profile as the origin, the airfoil chord as the x-axis, and the direction perpendicular to the chord as the z-axis, so as to obtain a preprocessed polyhedron model. The preprocessed polyhedron model is then identified by contour line interception and parameter extraction to obtain the airfoil model. In the contour line interception and parameter extraction, the parameters obtained include relative thickness, relative curvature and relative position of maximum curvature. Then, the key point constraints are introduced through the obtained airfoil model, the contour line is fitted to obtain the final contour line, and the blade model is reconstructed through the final contour line. On the other hand, a blade reverse modeling system based on airfoil parameter extraction is proposed. The system comprises a scanning module, a preprocessing module, an identification module, and a reverse reconstruction module. In the scanning module, a blade polyhedron model is obtained by scanning and extracting the blade. In the preprocessing module, the polyhedron model is padded and the coordinate system aligned to obtain a preprocessed polyhedron model. In the identification module, the final contour line is identified by contour interception and parameter extraction of the preprocessed polyhedron model. In the reverse modeling module, the blade model is reconstructed from the final contour line. At the ten-decimation points of the chord, perpendicular lines are drawn to intersect the contour line to obtain nine line segments. The midpoints of these nine line segments are connected sequentially to obtain the mean camber line of the blade airfoil. The mean camber line obtained by this method has a certain error. The error of the mean camber line at the ten-decimation points is small, but the error at other locations is random. If the error at individual ten-decimation points is large, it may lead to large errors in the maximum camber and the maximum camber position.

[0005] 2. CN106123725B - a method for inverse design of compressor blades to correct multi-dimensional processing errors, CN110781573A - a method for reverse design of aircraft engine rotors, CN114492062A - a method for reverse reconstruction of fan blades based on non-contact measurement. In the alignment stage, it is necessary to rely on the components or tenons of the blades to assist in positioning and alignment. The alignment method is relatively complicated, and the components and tenons of the blades are mostly used as the main load-bearing parts during operation. They are prone to wear or deformation after long-term use, which leads to alignment errors. In the reverse modeling stage, the contour line is extracted by curvature, and the blade is intercepted according to the direction of the contour line, which is also greatly affected by the wear or deformation of the blade. After the parameters are extracted, the existing methods mainly correct the parameters by removing the deviating points or calculating the average value through normal distribution. The model reconstructed using this data is closer to the model obtained directly by scanning. The forward design process of the blade is not taken into account, and there will be certain differences from the forward design CAD model.

[0006] 3. CN104008257A - A reverse design method for components with complex surfaces. It uses a large amount of data and has a complex calculation process. It also does not consider the forward design process of the blade and will have certain differences from the forward design CAD model. Summary of the Invention

[0007] The purpose of the present disclosure is to overcome the shortcomings of the existing technology and provide a NACA airfoil blade reverse modeling method, device, computer equipment and storage medium that effectively improves the accuracy of model recognition and reduces the data processing difficulty of airfoil blade reverse modeling.

[0008] The purpose of this disclosure is achieved through the following technical solutions:

[0009] A NACA airfoil blade reverse modeling method, the method comprising:

[0010] Scan the NACA airfoil blade to be identified to obtain a blade solid model;

[0011] Dividing the cross-sectional profile of the blade solid model into regions, and calculating corresponding airfoil shaping parameters according to the area values ​​of each region;

[0012] Outputting the model of the NACA airfoil blade to be identified according to the airfoil shaping parameters;

[0013] A lofting operation is performed on the cross-sectional profile according to the model to generate a NACA airfoil blade model.

[0014] In one embodiment, dividing the cross-sectional profile of the blade solid model into regions and calculating corresponding airfoil shaping parameters according to the area values ​​of each region includes: performing a cross-sectional operation on the blade solid model to obtain the airfoil profile.

[0015] In one embodiment, the section operation is performed on the blade solid model to obtain the airfoil profile, and then further includes: performing a profile curvature operation on the airfoil profile to obtain an airfoil chord distribution line; performing a perpendicular operation on the airfoil chord distribution line to obtain a chord perpendicular line passing through the upper and lower profile lines of the airfoil; and obtaining the four-image partition area of ​​the airfoil profile based on the airfoil chord distribution line and the chord perpendicular line.

[0016] In one embodiment, the performing a perpendicular operation on the airfoil chord distribution line includes: selecting a midpoint of the airfoil chord distribution line as a vertical point of the chord perpendicular line.

[0017] In one embodiment, the method of obtaining the four-image partition area of ​​the airfoil profile according to the airfoil chord distribution line and the chord perpendicular line further includes: performing a merging integral operation on the four-image partition area of ​​the airfoil profile to obtain the maximum thickness of the airfoil, and the calculation formula of the merging integral operation satisfies the following conditions: S 左上 +S 左下 +S 右上 +S 右下 =0.685t, where S 左上 is the area of ​​the upper left side of the four-image partition area of ​​the airfoil profile, S 左下 is the area of ​​the lower left side of the four-image partition area of ​​the airfoil profile, S 右上 is the area of ​​the upper right side of the four-image partition area of ​​the airfoil profile, S 右下 is the area of ​​the lower right side of the four-image partition area of ​​the airfoil profile, and t is the maximum thickness of the airfoil.

[0018] In one embodiment, the method of obtaining the four-image partition area of ​​the airfoil profile according to the airfoil chord distribution line and the chord perpendicular line further includes: performing a difference integral operation on the four-image partition area of ​​the airfoil profile to obtain the maximum camber of the airfoil, and the calculation formula of the difference integral operation satisfies the following conditions: Among them, S 左上 is the area of ​​the upper left side of the four-image partition area of ​​the airfoil profile, S 左下 is the area of ​​the lower left side of the four-image partition area of ​​the airfoil profile, S 右上 is the area of ​​the upper right side of the four-image partition area of ​​the airfoil profile, S 右下 is the area of ​​the lower right side of the four-image partition area of ​​the airfoil contour, and m is the maximum curvature of the airfoil.

[0019] In one embodiment, the method further comprises: performing a unilateral difference operation on the four-image partition area of ​​the airfoil profile according to the airfoil chord distribution line and the chord perpendicular line to obtain the maximum camber position of the airfoil, wherein the calculation formula of the unilateral difference operation satisfies the following conditions: Among them, S 左上 is the area of ​​the upper left side of the four-image partition area of ​​the airfoil profile, S 左下 is the area of ​​the lower left side of the four-image partition area of ​​the airfoil profile, S 右上 is the area of ​​the upper right side of the four-image partition area of ​​the airfoil profile, S 右下 is the area of ​​the lower right side of the four-image partition area of ​​the airfoil profile, and p is the position of the maximum curvature of the airfoil.

[0020] A NACA airfoil blade model identification device comprises: an airfoil blade acquisition module, an area shaping processing module, and an airfoil blade reverse generation module; the airfoil blade acquisition module is used to scan a NACA airfoil blade to be identified to obtain a blade entity model; the area shaping processing module is used to divide the cross-sectional profile of the blade entity model into regions, and calculate corresponding airfoil shaping parameters according to the area value of each region to obtain the airfoil shaping parameters; the airfoil blade reverse generation module is used to output the model of the NACA airfoil blade to be identified according to the airfoil shaping parameters, and perform a lofting operation on the cross-sectional profile according to the model to generate a NACA airfoil blade model.

[0021] A computer device includes a memory and a processor, wherein the memory stores a computer program, and when the processor executes the computer program, the following steps are implemented:

[0022] Scan the NACA airfoil blade to be identified to obtain a blade solid model;

[0023] Dividing the cross-sectional profile of the blade solid model into regions, and calculating corresponding airfoil shaping parameters according to the area values ​​of each region to obtain the airfoil shaping parameters;

[0024] Outputting the model of the NACA airfoil blade to be identified according to the airfoil shaping parameters;

[0025] A lofting operation is performed on the cross-sectional profile according to the model to generate a NACA airfoil blade model.

[0026] A computer-readable storage medium stores a computer program, which, when executed by a processor, implements the following steps:

[0027] Scan the NACA airfoil blade to be identified to obtain a blade solid model;

[0028] Dividing the cross-sectional profile of the blade solid model into regions, and calculating corresponding airfoil shaping parameters according to the area values ​​of each region to obtain the airfoil shaping parameters;

[0029] Outputting the model of the NACA airfoil blade to be identified according to the airfoil shaping parameters;

[0030] A lofting operation is performed on the cross-sectional profile according to the model to generate a NACA airfoil blade model.

[0031] Compared with the prior art, the present disclosure has at least the following advantages:

[0032] After scanning the NACA airfoil blade to be identified, its corresponding blade solid model is obtained. By using the area method to divide the blade solid model into regions, it is convenient to determine the regional area of ​​the NACA airfoil blade to be identified, so as to calculate the corresponding airfoil shaping parameters according to the regional area distribution. Finally, the model can be determined according to the airfoil shaping parameters, which reduces the error caused by the blade contour. Moreover, only the area of ​​the blade needs to be calculated, which reduces the data processing amount, effectively improves the model recognition accuracy and reduces the data processing difficulty of the reverse modeling of the airfoil blade. BRIEF DESCRIPTION OF THE DRAWINGS

[0033] In order to more clearly illustrate the technical solutions of the embodiments of the present disclosure, the following briefly introduces the drawings required for use in the embodiments. It should be understood that the following drawings only illustrate certain embodiments of the present disclosure and therefore should not be regarded as limiting the scope. For ordinary technicians in this field, other relevant drawings can be obtained based on these drawings without creative work.

[0034] Figure 1 is a flow chart of a method for reverse modeling of a NACA airfoil blade in one embodiment;

[0035] Figure 2 is a regional distribution diagram of an airfoil profile in one embodiment;

[0036] Figure 3 is a diagram of a dimensionless coordinate system of an airfoil in one embodiment;

[0037] Figure 4 is a regional distribution diagram of an airfoil profile in another embodiment;

[0038] Figure 5 A physical model diagram of an airfoil blade in one embodiment;

[0039] Figure 6 for Figure 5 A three-dimensional scan of the airfoil blade shown;

[0040] Figure 7for Figure 5 An airfoil cross-sectional profile of the airfoil blade shown;

[0041] Figure 8 for Figure 5 an area distribution diagram of the airfoil profile of the airfoil blade shown;

[0042] Figure 9 for Figure 5 Area plot of the area of ​​the blade of the shown airfoil;

[0043] Figure 10 for Figure 9 Area simulation diagram of the area shown;

[0044] Figure 11 for Figure 5 Forward modeling diagram of the airfoil blade shown;

[0045] Figure 12 for Figure 11 The body deviation analysis diagram of the forward modeled airfoil blade is shown;

[0046] Figure 13 A simulation diagram of an airfoil blade of an existing solution;

[0047] Figure 14 Another simulation diagram of an airfoil blade of the existing solution;

[0048] Figure 15 1 is a diagram of the internal structure of a computer device in one embodiment. DETAILED DESCRIPTION

[0049] To facilitate understanding of the present disclosure, a more comprehensive description of the present disclosure will be provided below with reference to the accompanying drawings. The accompanying drawings illustrate preferred embodiments of the present disclosure. However, the present disclosure can be implemented in many different forms and is not limited to the embodiments described herein. Rather, these embodiments are provided to provide a more thorough and comprehensive understanding of the disclosure.

[0050] It should be noted that when an element is referred to as being "fixed to" another element, it may be directly attached to the other element or there may be an intermediate element. When an element is referred to as being "connected to" another element, it may be directly connected to the other element or there may be an intermediate element. The terms "vertical," "horizontal," "left," "right," and similar expressions used herein are for illustrative purposes only and do not represent the only implementation methods.

[0051] Unless otherwise defined, all technical and scientific terms used herein have the same meaning as commonly understood by those skilled in the art to which this disclosure pertains. The terms used herein in the specification of this disclosure are intended only to describe specific embodiments and are not intended to limit this disclosure. The term "and / or" as used herein includes any and all combinations of one or more of the associated listed items.

[0052] The present disclosure relates to a method for reverse modeling a NACA airfoil blade. In one embodiment, the method comprises scanning a NACA airfoil blade to be identified to obtain a blade solid model; dividing the cross-sectional profile of the blade solid model into regions, and calculating corresponding airfoil shaping parameters based on the area values ​​of each region to obtain the airfoil shaping parameters; outputting a model of the NACA airfoil blade to be identified based on the airfoil shaping parameters; and performing a lofting operation on the cross-sectional profile based on the model to generate a NACA airfoil blade model. After scanning the NACA airfoil blade to be identified, a corresponding blade solid model is obtained. The blade solid model is divided into regions using an area method to facilitate determining the area of ​​the NACA airfoil blade to be identified, thereby facilitating calculation of corresponding airfoil shaping parameters based on the regional area distribution. Finally, the model is determined based on the airfoil shaping parameters, reducing errors caused by the blade profile. Furthermore, only the blade area needs to be calculated, which reduces the amount of data processing, effectively improving the accuracy of model identification, and reducing the data processing difficulty of reverse modeling of the airfoil blade.

[0053] See also Figure 1 , which is a flow chart of a NACA airfoil blade reverse modeling method according to an embodiment of the present disclosure. The NACA airfoil blade reverse modeling method includes some or all of the following steps.

[0054] S100: Scanning the NACA airfoil blade to be identified to obtain a blade entity model.

[0055] In this embodiment, the NACA airfoil blade to be identified is a blade that requires model identification of the airfoil blade. The model of the NACA airfoil blade to be identified serves as a prerequisite for the reverse reconstruction of the NACA airfoil blade, that is, it plays an important role in repairing and reconstructing the NACA airfoil blade to be identified and forming an analytical design. The blade scanning of the NACA airfoil blade to be identified is to measure the contour data of the three-dimensional surface of the NACA airfoil blade to be identified, for example, by contact measurement and non-contact measurement. The present disclosure adopts a non-contact data measurement method, applies three-dimensional digital laser scanning measurement technology, collects a point cloud of the blade solid surface, and pre-processes the data point cloud. With the help of three-dimensional solid modeling software and related technologies, the blade solid model is reconstructed, thereby facilitating the determination of the blade solid model.

[0056] S200: Dividing the cross-sectional profile of the blade solid model into regions, and calculating corresponding airfoil shaping parameters according to the area values ​​of the respective regions to obtain the airfoil shaping parameters.

[0057] In this embodiment, the blade solid model is the three-dimensional object contour data of the NACA airfoil blade to be identified. The blade solid model serves as the surface contour of the NACA airfoil blade to be identified, facilitating determination of the three-dimensional surface shaping of the NACA airfoil blade to be identified. The cross-sectional contour of the blade solid model is divided into regions, and corresponding airfoil shaping parameters are calculated based on the area values ​​of each region. This involves using an area method to divide the cross-sectional contour of the blade solid model. Specifically, the cross-sectional contour of the blade solid model is divided into areas, and corresponding airfoil shaping parameters are calculated based on the divided areas. The airfoil shaping parameters correspond to the contour features of the NACA airfoil blade to be identified, facilitating determination of the airfoil characteristics of the NACA airfoil blade to be identified.

[0058] S300: Outputting the model of the NACA airfoil blade to be identified according to the airfoil shaping parameters.

[0059] In this embodiment, the airfoil shaping parameters are obtained based on the blade solid model, which is the three-dimensional object contour data of the NACA airfoil blade to be identified. The blade solid model serves as the surface contour of the NACA airfoil blade to be identified, facilitating determination of the three-dimensional surface shaping of the NACA airfoil blade to be identified. The cross-sectional contour of the blade solid model is divided into regions, and corresponding airfoil shaping parameters are calculated based on the area values ​​of each region. This involves using an area method to divide the cross-sectional contour of the blade solid model. Specifically, the cross-sectional contour of the blade solid model is divided into areas, and corresponding airfoil shaping parameters are calculated based on the divided areas. The airfoil shaping parameters correspond to the contour features of the NACA airfoil blade to be identified, facilitating determination of the airfoil characteristics of the NACA airfoil blade to be identified. The airfoil shaping parameters serve as a standard for distinguishing airfoil blade models. For example, the airfoil shaping parameters include at least one of the maximum thickness of the airfoil, the maximum camber of the airfoil, and the position of the maximum camber of the airfoil. By characterizing the airfoil shaping parameters, it is convenient to quickly determine the specific model of the NACA airfoil blade to be identified.

[0060] S400: performing a lofting operation on the cross-sectional profile according to the model to generate a NACA airfoil blade model.

[0061] In this embodiment, after scanning the NACA airfoil blade to be identified, its corresponding blade solid model is obtained. By adopting the area method to divide the blade solid model into regions, it is convenient to determine the regional area of ​​the NACA airfoil blade to be identified, thereby facilitating the calculation of the corresponding airfoil shaping parameters based on the regional area distribution. Finally, the model can be determined based on the airfoil shaping parameters, thereby reducing the error caused by the blade contour. Moreover, only the area of ​​the blade needs to be calculated, which reduces the data processing amount, effectively improves the model identification accuracy, and reduces the data processing difficulty of the reverse modeling of the airfoil blade.

[0062] In one embodiment, dividing the cross-sectional profile of the blade solid model into regions and calculating corresponding airfoil shaping parameters based on the area values ​​of each region to obtain the airfoil shaping parameters includes: performing a cross-sectional operation on the blade solid model to obtain the airfoil profile. In this embodiment, the blade solid model is the three-dimensional object contour data of the NACA airfoil blade to be identified. The blade solid model serves as the surface contour of the NACA airfoil blade to be identified, facilitating determination of the three-dimensional surface shaping of the NACA airfoil blade to be identified. Dividing the cross-sectional profile of the blade solid model into regions and calculating corresponding airfoil shaping parameters based on the area values ​​of each region involves dividing the cross-sectional profile of the blade solid model using an area method. Specifically, the cross-sectional profile of the blade solid model is divided into areas, and corresponding airfoil shaping parameters are calculated based on the divided areas. The airfoil shaping parameters correspond to the contour features of the NACA airfoil blade to be identified, facilitating determination of the airfoil characteristics of the NACA airfoil blade to be identified. In the process of dividing the cross-sectional profile into regions and calculating the corresponding airfoil shaping parameters according to the area values ​​of each region, a cross-sectional operation is first performed on the blade solid model so as to slice the blade solid model to obtain the precise airfoil profile of the NACA airfoil blade to be identified. Specifically, the blade solid model is face-aligned, that is, the plane at the top surface of the polyhedron is used as the reference plane for alignment. After the alignment is completed, the blade model is cross-sectionally acquired using an equidistant interception method to obtain the airfoil profile.

[0063] Furthermore, performing a cross-sectional operation on the solid blade model to obtain an airfoil profile further includes: performing a profile curvature operation on the airfoil profile to obtain an airfoil chord distribution line; performing a perpendicular operation on the airfoil chord distribution line to obtain a chord perpendicular line passing through the upper and lower profile lines of the airfoil; and obtaining four-dimensional partitioned areas of the airfoil profile based on the airfoil chord distribution line and the chord perpendicular line. In this embodiment, the solid blade model serves as the surface profile of the NACA airfoil blade to be identified, facilitating determination of the three-dimensional surface shaping of the NACA airfoil blade to be identified. The cross-sectional profile of the solid blade model is partitioned into regions, and corresponding airfoil shaping parameters are calculated based on the area values ​​of each region. Specifically, the cross-sectional profile of the solid blade model is partitioned into regions, and corresponding airfoil shaping parameters are calculated based on the partitioned areas. The airfoil shaping parameters correspond to the profile features of the NACA airfoil blade to be identified, facilitating determination of the airfoil characteristics of the NACA airfoil blade to be identified. During the process of dividing the cross-sectional profile into regions and calculating the corresponding airfoil shaping parameters based on the area values ​​of each region, a cross-sectional operation is first performed on the blade solid model to facilitate slicing of the blade solid model to obtain the precise airfoil profile of the NACA airfoil blade to be identified. Specifically, the blade solid model is face-aligned, i.e., alignment is performed using the plane at the top surface of the polyhedron as the reference plane. After alignment is completed, a cross-sectional capture of the blade model is performed using an equidistant interception method to obtain the airfoil profile. The airfoil profile corresponds to the precise airfoil profile of the NACA airfoil blade to be identified. By operating on the profile curvature of the airfoil profile, the chord line of the airfoil blade is determined based on the profile curvature distribution of the NACA airfoil blade to be identified. Specifically, based on the curvature calculation, the point with the maximum rear end curvature is found, i.e., the tail of the airfoil, which is taken as the trailing edge B. A circle tangent to the front end of the airfoil is drawn with the trailing edge point B as the center, and the center of the circle is connected to the tangent point. Since the radius is the longest line segment passing through the center of the semicircle, the tangent point is the point where the distance between the front and trailing edges of the airfoil is the largest, that is, the leading edge A of the airfoil, and the straight line connecting them is the chord.

[0064] After determining the chord distribution line, i.e., the chord of the airfoil blade, a perpendicular to the chord distribution line is manipulated to create a corresponding chord perpendicular using the chord as a reference. Furthermore, this chord perpendicular passes through the upper and lower contour lines of the airfoil profile. For example, the midpoint of the chord distribution line is selected as the perpendicular point of the chord perpendicular. Thus, after the chord line and the chord perpendicular are determined, the chord line and the chord perpendicular divide the airfoil profile into regions to form four-image partition areas of the airfoil profile. Specifically, the four-image partition areas of the airfoil profile are the four divided regions formed on the airfoil profile, thereby facilitating area calculation for each region on the airfoil profile.

[0065] Furthermore, the method further comprises: performing a merging integral operation on the four-image partition areas of the airfoil profile according to the airfoil chord distribution line and the chord perpendicular line to obtain the maximum thickness of the airfoil, wherein the calculation formula of the merging integral operation satisfies the following conditions: S 左上 +S 左下 +S 右上 +S 右下 =0.685t, where S 左上 is the area of ​​the upper left side of the four-image partition area of ​​the airfoil profile, S 左下 is the area of ​​the lower left side of the four-image partition area of ​​the airfoil profile, S 右上 is the area of ​​the upper right side of the four-image partition area of ​​the airfoil profile, S 右下 is the area of ​​the lower right side of the four-image partition area of ​​the airfoil profile, and t is the maximum thickness of the airfoil.

[0066] In this embodiment, the four-image partition area of ​​the airfoil profile is obtained based on the airfoil profile, and is obtained after the airfoil profile is divided by chord and vertical lines. The four-image partition area of ​​the airfoil profile is used as the area of ​​the four divided regions of the airfoil profile. Figure 2 .

[0067] The acquisition of the four-image partition area of ​​the airfoil profile is based on the dimensionless coordinate system of the airfoil, such as Figure 3 As shown in the dimensionless coordinate system of the airfoil, the leading edge point A is the origin, the chord direction is the x-axis, the direction perpendicular to the chord is the y-axis, and all length values ​​are divided by the chord length c. Therefore, in this coordinate system, the coordinate of the trailing edge B is (1, 0). c is the curvature value d of the change in x, and the corresponding function is the curvature distribution function, y t It is 1 / 2 of the thickness value h with respect to the change of x. The corresponding function is the thickness distribution function. The expressions of the two distribution functions of the NACA four-digit airfoil are:

[0068]

[0069] In these two distribution functions, the parameter m is the maximum curvature, the parameter p is the position of the maximum curvature, and the parameter t is the maximum thickness. Since the value of the chord length c in the dimensionless coordinate system is 1, the value of the maximum relative curvature is also m, the value of the relative position of the maximum curvature is also p, and the value of the maximum relative thickness is also t.

[0070] The upper and lower profiles of the airfoil can be expressed as the superposition of two distribution functions, namely:

[0071] y u =yc +y t (3)

[0072] y l =y c -y t (4)

[0073] y u is the y-coordinate value of the upper profile line of the airfoil, l The four areas corresponding to the four-image partition area of ​​the airfoil profile are S 左上 、S 左下 、S 右上 、S 右下 .

[0074] The total area of ​​the NACA airfoil to be identified is calculated using the integral formula and the upper and lower profile functions, namely:

[0075]

[0076] Substituting the curvature distribution function and thickness distribution function into the upper and lower profile functions, we can obtain:

[0077]

[0078] Integrating this formula, we get:

[0079] S 左上 +S 左下 +S 右上 +S 右下 =0.685t (7)

[0080] Therefore, an area formula for the parameter t can be obtained, which makes it easy to obtain the maximum thickness t of the airfoil by combining and integrating the areas of the four partitions of the airfoil profile, thereby facilitating the determination of a key characteristic parameter of the NACA airfoil blade to be identified in the model identification process.

[0081] In another embodiment, the step of obtaining the four-image partition areas of the airfoil profile according to the airfoil chord distribution line and the chord perpendicular line further includes: performing a difference integral operation on the four-image partition areas of the airfoil profile to obtain the maximum camber of the airfoil, wherein the calculation formula of the difference integral operation satisfies the following conditions: Among them, S 左上 is the area of ​​the upper left side of the four-image partition area of ​​the airfoil profile, S 左下 is the area of ​​the lower left side of the four-image partition area of ​​the airfoil profile, S 右上 is the area of ​​the upper right side of the four-image partition area of ​​the airfoil profile, S 右下is the area of ​​the lower right side of the four-image partition area of ​​the airfoil contour, and m is the maximum curvature of the airfoil.

[0082] In this embodiment, in addition to the chord line, the airfoil profile also has a camber line, which is a curve connecting the leading and trailing edges. In the normal direction of the chord line, the distances from the points on the curve to the upper and lower surfaces of the airfoil are equal. The airfoil profile is divided into three area regions according to the distribution of the chord line and the camber line. Figure 4 The areas corresponding to the three regions are S1, S2, and S3, and the corresponding area formulas are:

[0083]

[0084] Through these three area formulas, we can get:

[0085]

[0086] Where (S1+S2) represents the area above the chord, which can be obtained by S 左上 and S 右上 The sum of S3 represents the area below the chord, which can be expressed by S 左下 and S 右下 The sum of , that is:

[0087]

[0088] Integrating this formula, we get:

[0089]

[0090] Therefore, an area formula for the parameter m can be obtained, which makes it easy to obtain the maximum curvature m of the airfoil by combining and integrating the areas of the four partitions of the airfoil profile, thereby facilitating the determination of another key characteristic parameter of the NACA airfoil blade to be identified in the model identification process.

[0091] In another embodiment, the method of obtaining the four-image partition area of ​​the airfoil profile according to the airfoil chord distribution line and the chord perpendicular line further includes: performing a unilateral difference operation on the four-image partition area of ​​the airfoil profile to obtain the maximum camber position of the airfoil, and the calculation formula of the unilateral difference operation satisfies the following conditions:

[0092]

[0093] Among them, S 左上 is the area of ​​the upper left side of the four-image partition area of ​​the airfoil profile, S 左下 is the area of ​​the lower left side of the four-image partition area of ​​the airfoil profile, S 右上is the area of ​​the upper right side of the four-image partition area of ​​the airfoil profile, S 右下 is the area of ​​the lower right side of the four-image partition area of ​​the airfoil profile, and p is the position of the maximum curvature of the airfoil.

[0094] In this embodiment, according to formula (11), we can obtain:

[0095]

[0096] Because y c It is a piecewise function. Using either formula (14) or formula (15) alone can obtain the area formula for the parameter p. However, the selection of the maximum curvature position p of the airfoil is segmented. Therefore, we consider each case separately and use the two formulas at the same time. When p < 0.5, the corresponding parameter p is obtained by formula (15). When p ≥ 0.5, the corresponding parameter p is obtained by formula (14). The final area formula is:

[0097]

[0098] Substituting the value of m into equations (16) and (17), we can obtain an area formula for the parameter p, which is convenient for obtaining the maximum curvature position p of the airfoil by integrating the areas of the four partitions of the airfoil profile, thereby facilitating the determination of another key characteristic parameter of the NACA airfoil blade to be identified in the model identification process. Therefore, it is only necessary to obtain S 左上 、S 左下 、S 右上 、S 右下 The numerical values ​​of the four areas can be directly calculated by the four formulas (7), (13), (16) and (17) to obtain the parameters t, m and p. When obtaining the area of ​​the selected area, the area value can be automatically obtained by the area calculation function provided in the commonly used CAD software, such as Solidworks, NX, Geomagic Design X, etc. In another embodiment, the area calculation function usually measures the specified surface and does not directly measure the area enclosed by the specified curve. Therefore, it is necessary to generate the surface enclosed by the curve through the "stretch" function or the "surface generation" function to measure it. Among them, the equation after substituting the m value is a quadratic equation. According to the properties of the quadratic equation, it can be seen that the equation has two solutions. Since the size of the p value is unknown, it is necessary to substitute the area into the two equations of the piecewise function respectively, so that four solutions will be obtained. The p value must be a real number and within the range of values ​​specified by the piecewise function. The values ​​that do not meet the requirements must be discarded, thereby obtaining the correct parameter p.

[0099] In another embodiment, the NACA airfoil blade to be identified is scanned to obtain a blade solid model, and then further includes: performing facet preprocessing on the blade solid model. Specifically, the polyhedron model obtained by scanning, i.e., the blade solid model, is imported into the software, and the face filling function is used to repair the missing parts of the edge of the blade polyhedron model. Since the curvature of the curved surface edge varies greatly, a bridging filling method is selected. After the filling is completed, if the edge is not smooth and there is a large error with the actual object, the filled part is smoothed to improve the accuracy of the subsequent regional division of the cross-sectional contour, thereby facilitating the subsequent acquisition of the accurate area of ​​each region, so that the recognition accuracy of the model of the NACA airfoil blade to be identified is further improved.

[0100] In one embodiment, the specific operations of the NACA airfoil blade reverse modeling method are as follows:

[0101] Step 1. 3D scanning of NACA airfoil blades to be identified. For details on the blade model, see Figure 5 .

[0102] Step 2. Pre-process the surface file of the blade solid model, that is, the surface file after 3D scanning and defect processing. See the attached Figure 6 .

[0103] Step 3. Extract the airfoil section profile from the patch file, i.e. the airfoil profile. Figure 7 .

[0104] Step 4. Find the chord line through curvature analysis and tangent point, draw the perpendicular midline of the chord line, and divide the airfoil profile into 4 areas. See the attached diagram for details. Figure 8 .

[0105] Step 5. Convert each area into a surface and obtain the corresponding area value based on the current surface. See the attached Figure 9 .

[0106] Step 6. Calculate the maximum thickness t, maximum curvature m, and maximum curvature position p of the airfoil according to the airfoil area formula. The areas of the four regions are measured as S 左上 =50.5563mm 2 、S 左下 =24.2427mm 2 、S 右上 =32.7198mm 2 、S 右下 =11.3883mm 2 , see Figure 10Since the design equation for deriving the airfoil area formula is a dimensionless equation, its values ​​are normalized, that is, they are divided by the chord length c, so the obtained area also needs to be normalized, that is, they are divided by the chord length c. 2 , that is, S 左上 =0.029043, S 右上 =0.018826, S 左下 =0.013994, S 右下 =0.006552.

[0107] Substituting the area values ​​into the airfoil area formula with parameters t and m yields t = 0.099876 and m = 0.020492. Rounding t and m yields t = 0.10 and m = 0.02. Substituting the area values ​​and m into the airfoil area formula with parameter p yields p1 = 0.766288, p2 = 0.418982, p3 = 0.332248 + 0.018990i, and p4 = 0.332248 - 0.018990i. Since p should be within the range and a real number, p = p2 = 0.418982. Rounding p yields p = 0.4, indicating that the NACA airfoil blade to be identified is NACA2410.

[0108] Step 7. According to the model of the NACA airfoil blade to be identified, forward modeling is performed on it, that is, the cross-sectional profile is lofted according to the model to generate a NACA airfoil blade model, such as Figure 11 The lofting operation for the cross-sectional profile is to substitute the model into the above airfoil formula so as to inversely determine the cross-sectional shape of the NACA airfoil blade.

[0109] Step 8. Perform body deviation analysis on the forward modeling. Through the analysis diagram, it can be found that the green part occupies a larger area, and the deviation of other parts is mainly yellow. There is no red with large deviation, which meets the tolerance requirements of the blade. Figure 12 shown.

[0110] In one embodiment, the NACA airfoil blade reverse modeling method involved in the present disclosure has airfoil parameter calculation differences from the existing blade reverse modeling method based on airfoil parameter extraction, as follows:

[0111] For the same NACA airfoil blade, the actual airfoil parameters are m = 0.04, p = 0.4. The existing solution draws the median arc line through the midpoint of the ten-equal points, such as Figure 13 As shown, the maximum distance point between the mid-arc and the chord is found through the tangent relationship, as shown in Figure 14 As shown, finally passing through the tangent point, the software measured m = 0.019515, p = 0.417468.

[0112] The present disclosure finds the chord through curvature analysis and tangent points, draws the perpendicular bisector of the chord, measures the areas of the four regions, and then plugs them into the area formula to obtain m=0.019552 and p=0.385927.

[0113] It can be seen from this that the existing solution has Δm=0.000485, Δp=0.017468, and the present disclosure has Δm=0.000448, which is 0.000037 smaller than the existing solution, and Δp=0.014073, which is 0.003395 smaller than the existing solution.

[0114] Therefore, it can be found that the deviations of the two schemes are very close and very small, but separately, the deviation of the present invention will be smaller than the deviation of the existing scheme, and the auxiliary elements involved include the median arc line and the median perpendicular line of the median arc line. In the acquisition of the airfoil section, only surface operations need to be performed to intercept it, which is easier to operate.

[0115] In one embodiment, the present disclosure also relates to a NACA airfoil blade model identification device, comprising: an airfoil blade acquisition module, an area shaping processing module and an airfoil blade inverse generation module; the airfoil blade acquisition module is used to perform blade scanning on the NACA airfoil blade to be identified to obtain a blade solid model; the area shaping processing module is used to divide the cross-sectional contour of the blade solid model into regions, and calculate the corresponding airfoil shaping parameters according to the area value of each region to obtain the airfoil shaping parameters; the airfoil blade inverse generation module is used to output the model of the NACA airfoil blade to be identified according to the airfoil shaping parameters, and perform lofting operations on the cross-sectional contour according to the model to generate a NACA airfoil blade model.

[0116] In this embodiment, after the airfoil blade acquisition module scans the NACA airfoil blade to be identified, the corresponding blade entity model is obtained. The area shaping processing module divides the blade entity model into regions by using the area method, so as to facilitate the determination of the regional area of ​​the NACA airfoil blade to be identified, thereby facilitating the calculation of the corresponding airfoil shaping parameters based on the regional area distribution. Finally, the airfoil blade reverse generation module can determine the model based on the airfoil shaping parameters, thereby reducing the error caused by the blade contour. Moreover, only the area of ​​the blade needs to be calculated, which reduces the amount of data processing, effectively improves the accuracy of model recognition, and reduces the data processing difficulty of airfoil blade reverse modeling.

[0117] In one embodiment, a computer device is provided. The computer device may be a server, and its internal structure diagram may be as follows: Figure 15As shown. The computer device includes a processor, a memory and a network interface connected via a system bus. The processor of the computer device is used to provide computing and control capabilities. The memory of the computer device includes a non-volatile storage medium and an internal memory. The non-volatile storage medium stores an operating system, a computer program and a database. The internal memory provides an environment for the operation of the operating system and the computer program in the non-volatile storage medium. The database of the computer device is used to store data such as blade solid models, airfoil shaping parameters and models of NACA airfoil blades to be identified. The network interface of the computer device is used to communicate with an external terminal via a network connection. When the computer program is executed by the processor, a NACA airfoil blade inverse modeling method is implemented.

[0118] Those skilled in the art will understand that Figure 15 The structure shown in the figure is only a block diagram of a part of the structure related to the solution of the present application, and does not constitute a limitation on the computer device to which the solution of the present application is applied. The specific computer device may include more or fewer components than shown in the figure, or combine certain components, or have a different component arrangement.

[0119] In one embodiment, the present application further provides a computer device, including a memory and a processor, wherein a computer program is stored in the memory, and the processor implements the steps in the above-mentioned method embodiments when executing the computer program.

[0120] In one embodiment, the present application further provides a computer-readable storage medium having a computer program stored thereon, which implements the steps in the above-mentioned method embodiments when the computer program is executed by a processor.

[0121] Those skilled in the art will appreciate that all or part of the processes in the above-mentioned embodiment methods can be implemented by instructing the relevant hardware through a computer program, and the computer program can be stored in a non-volatile computer-readable storage medium. When the computer program is executed, it can include the processes of the embodiments of the above-mentioned methods. Among them, any reference to memory, storage, database or other media used in the embodiments provided in this application may include at least one of non-volatile and volatile memory. Non-volatile memory may include read-only memory (ROM), magnetic tape, floppy disk, flash memory or optical memory, etc. Volatile memory may include random access memory (RAM) or external cache memory. As an illustration and not limitation, RAM can be in various forms, such as static random access memory (SRAM) or dynamic random access memory (DRAM).

[0122] The above-described embodiments merely represent several implementation methods of the present disclosure. While the descriptions are relatively specific and detailed, they should not be construed as limiting the scope of the present invention. It should be noted that a person of ordinary skill in the art could make various modifications and improvements without departing from the spirit of the present disclosure, all of which fall within the scope of protection of the present disclosure. Therefore, the scope of protection of the present patent shall be determined by the appended claims.

Claims

1. A NACA airfoil blade reverse modeling method, characterized in that: include: Scan the NACA airfoil blade to be identified to obtain a blade solid model; Dividing the cross-sectional profile of the blade solid model into regions, and calculating corresponding airfoil shaping parameters according to the area values ​​of each region; Outputting the model of the NACA airfoil blade to be identified according to the airfoil shaping parameters; performing a lofting operation on the cross-sectional profile according to the model to generate a NACA airfoil blade model; The step of dividing the cross-sectional profile of the blade solid model into regions and calculating corresponding airfoil shaping parameters according to the area values ​​of each region includes: Performing a cross-section operation on the blade solid model to obtain an airfoil profile; The step of performing a cross-section operation on the blade solid model to obtain an airfoil profile further includes: Performing a profile curvature operation on the airfoil profile to obtain an airfoil chord distribution line; Performing a perpendicular operation on the airfoil chord distribution line to obtain a chord perpendicular line passing through the upper and lower profile lines of the airfoil; Obtaining the four-image partition area of ​​the airfoil profile according to the airfoil chord distribution line and the chord perpendicular line; The method further comprises: obtaining the four-image partition area of ​​the airfoil profile according to the airfoil chord distribution line and the chord perpendicular line; A combined integral operation is performed on the four-image partition area of ​​the airfoil profile to obtain the maximum thickness of the airfoil. The calculation formula of the combined integral operation meets the following conditions: , where S 左上 is the area of ​​the upper left side of the four-image partition area of ​​the airfoil profile, S 左下 is the area of ​​the lower left side of the four-image partition area of ​​the airfoil profile, S 右上 is the area of ​​the upper right side of the four-image partition area of ​​the airfoil profile, S 右下 is the area of ​​the lower right side of the four-image partition area of ​​the airfoil profile, and t is the maximum thickness of the airfoil.

2. The NACA airfoil blade reverse modeling method according to claim 1, characterized in that: The performing a perpendicular operation on the airfoil chord distribution line includes: The midpoint of the airfoil chord distribution line is selected as the vertical point of the chord perpendicular line.

3. The NACA airfoil blade reverse modeling method according to claim 1, characterized in that: The method further comprises: obtaining the four-image partition area of ​​the airfoil profile according to the airfoil chord distribution line and the chord perpendicular line; Perform a difference integration operation on the four-image partition area of ​​the airfoil profile to obtain the maximum camber of the airfoil. The calculation formula of the difference integration operation satisfies the following conditions: , where S 左上 is the area of ​​the upper left side of the four-image partition area of ​​the airfoil profile, S 左下 is the area of ​​the lower left side of the four-image partition area of ​​the airfoil profile, S 右上 is the area of ​​the upper right side of the four-image partition area of ​​the airfoil profile, S 右下 is the area of ​​the lower right side of the four-image partition area of ​​the airfoil contour, and m is the maximum curvature of the airfoil.

4. The NACA airfoil blade reverse modeling method according to claim 3, characterized in that: The method further comprises: obtaining the four-image partition area of ​​the airfoil profile according to the airfoil chord distribution line and the chord perpendicular line; A unilateral difference operation is performed on the four-image partition area of ​​the airfoil profile to obtain the maximum camber position of the airfoil. The calculation formula of the unilateral difference operation meets the following conditions: ； Among them, S 左上 is the area of ​​the upper left side of the four-image partition area of ​​the airfoil profile, S 左下 is the area of ​​the lower left side of the four-image partition area of ​​the airfoil profile, S 右上 is the area of ​​the upper right side of the four-image partition area of ​​the airfoil profile, S 右下 is the area of ​​the lower right side of the four-image partition area of ​​the airfoil profile, and p is the position of the maximum curvature of the airfoil.

5. A NACA airfoil blade model identification device using the NACA airfoil blade reverse modeling method according to any one of claims 1 to 4, characterized in that: include: An airfoil blade acquisition module, wherein the airfoil blade acquisition module is used to scan the NACA airfoil blade to be identified to obtain a blade entity model; An area shaping processing module, the area shaping processing module is used to divide the cross-sectional contour of the blade solid model into regions, and calculate corresponding airfoil shaping parameters according to the area value of each region; The airfoil blade reverse generation module is used to output the model of the NACA airfoil blade to be identified according to the airfoil shaping parameters, and to perform a lofting operation on the cross-sectional profile according to the model to generate a NACA airfoil blade model.

6. A computer device comprising a memory and a processor, wherein the memory stores a computer program, wherein: When the processor executes the computer program, the steps of the method according to any one of claims 1 to 4 are implemented.

7. A computer-readable storage medium having a computer program stored thereon, characterized in that: When the computer program is executed by a processor, the steps of the method according to any one of claims 1 to 4 are implemented.

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