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

Abstract

The invention provides an NACA airfoil blade reverse modeling method and device, computer equipment and a storage medium. The method comprises the following steps: performing blade scanning on a to-be-identified NACA airfoil blade to obtain a blade entity model; carrying out section contour region division on the blade entity model, and calculating corresponding airfoil shaping parameters according to the area value of each region; outputting the model of the to-be-identified NACA airfoil blade according to the airfoil shaping parameters; and carrying out lofting operation on the section contour according to the model so as to generate an NACA airfoil blade model. Regional area division is performed on the blade entity model by adopting an area method, so that the corresponding airfoil shaping parameters can be conveniently calculated according to regional area distribution, finally, the model can be determined according to the airfoil shaping parameters, errors caused by the blade contour are reduced, only the area of the blade needs to be calculated, the data processing amount is reduced, and the data processing efficiency is improved. The model identification accuracy is effectively improved, and the data processing difficulty of airfoil blade reverse modeling is reduced.

Description

Technical Field

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

[0002] Blades are key components of wind turbines and water turbines. NACA airfoil blades are widely used in both mechanical structures due to their special structure. While the blades are widely used, blade damage caused by cavitation and cavitation is gradually increasing. To address this problem, these blades need to be reverse modeled and redesigned. In addition, reverse modeling of NACA airfoil blades helps 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 method and system for reverse modeling of a blade based on airfoil parameter extraction. On the one hand, a method for reverse modeling of a blade is proposed. The blade is first scanned to obtain a polyhedron model of the blade, and then the missing parts of the obtained model are filled, and after filling, the filled area is smoothed to obtain a smoothed model. The top surface plane of the smoothed model is used as the "xoz" plane, the tangent point of the leading edge of the airfoil contour is used as the origin, the airfoil chord is used as the x-axis, and the direction perpendicular to the chord is used as the z-axis. The smoothed model is aligned with the coordinate system to obtain a preprocessed polyhedron model, and then the preprocessed polyhedron model is 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, which includes 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 preprocessed polyhedron model is obtained by filling the polyhedron model and aligning the coordinate system. 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 by the final contour line. At the ten-equal points of the chord, a vertical line is drawn to intersect the contour line to obtain 9 line segments, and the midpoints of the 9 line segments are connected in sequence to obtain the mid-arc line of the blade airfoil. The mid-arc line obtained by this method has a certain error. The error of the mid-arc line at the ten-equal points is small, but the error at other positions is random. If the error of individual ten-equal points is large, it may lead to large errors in the maximum curvature and the maximum curvature position obtained.

[0005] 2. CN106123725B-An inverse method for correcting multi-dimensional processing errors of compressor blades, 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 force-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 method mainly corrects the parameters by removing the deviation points or finding the average value through the normal distribution. The model reconstructed using this data is closer to the model directly obtained by scanning. The forward design process of the blade is not considered, and there will be some differences from the CAD model of the forward design.

[0006] 3. CN104008257A - A reverse design method for components with complex curved 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 be somewhat different from the forward design CAD model. Summary of the invention

[0007] The purpose of the present invention is to overcome the deficiencies in the prior art and to provide a NACA airfoil blade reverse modeling method, apparatus, computer equipment and storage medium that effectively improves the accuracy of model recognition and reduces the difficulty of data processing for reverse modeling of airfoil blades.

[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, the dividing of the cross-sectional contour of the blade solid model into regions and calculating the corresponding airfoil shaping parameters according to the area value of each region includes: performing cross-sectional operations on the blade solid model to obtain the airfoil contour.

[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 vertical line operation on the airfoil chord distribution line to obtain a chord vertical 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 vertical line.

[0016] In one of the embodiments, 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 four-image partition area of ​​the airfoil profile is obtained according to the airfoil chord distribution line and the chord perpendicular line, and then the four-image partition area of ​​the airfoil profile is combined and integrated to obtain the maximum thickness of the airfoil, and the calculation formula of the combined integration 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 area in 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 It is the area of ​​the lower right side of the four-image partition area of ​​the airfoil contour, and t is the maximum thickness of the airfoil.

[0018] In one embodiment, the step 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 area in 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 It 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 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: Among them, S 左上 is the area of ​​the upper left area in 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 It is the area of ​​the lower right side of the four-image partition area of ​​the airfoil contour, 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 perform blade scanning on a NACA airfoil blade to be identified to obtain a blade entity model; the area shaping processing module is used to perform regional division of the cross-sectional profile of the blade entity model, 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 lofting operation on the cross-sectional profile according to the model to generate a NACA airfoil blade model.

[0021] A computer device comprises 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 invention has at least the following advantages:

[0032] After scanning the NACA airfoil blade to be identified, the corresponding blade solid model is obtained. The area method is used to divide the blade solid model into regions, which facilitates 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 model can be determined based on the airfoil shaping parameters, 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 improving the accuracy of model recognition and reducing the data processing difficulty of reverse modeling of airfoil blades. BRIEF DESCRIPTION OF THE DRAWINGS

[0033] In order to more clearly illustrate the technical solutions of the embodiments of the present disclosure, the drawings required for use in the embodiments will be briefly introduced below. It should be understood that the following drawings only show certain embodiments of the present disclosure and therefore should not be regarded as limiting the scope. For ordinary technicians in this field, other related drawings can be obtained based on these drawings without paying 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 an 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] Fig. 9 for Figure 5 Area plot of the area of ​​the airfoil blade shown;

[0043] Fig.10 for Fig. 9 Area simulation diagram of the area of ​​the region shown;

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

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

[0046] Fig.13 A simulation schematic diagram of an airfoil blade of an existing solution;

[0047] Fig.14 Another simulation schematic diagram of an airfoil blade of a conventional solution;

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

[0049] In order to facilitate the understanding of the present disclosure, the present disclosure will be described more fully below with reference to the relevant drawings. The preferred embodiments of the present disclosure are given in the drawings. However, the present disclosure can be implemented in many different forms and is not limited to the embodiments described herein. On the contrary, the purpose of providing these embodiments is to make the disclosure of the present disclosure more thoroughly and comprehensively understood.

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

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

[0052] The present disclosure relates to a NACA airfoil blade reverse modeling method. In one embodiment, the NACA airfoil blade reverse modeling method includes scanning a NACA airfoil blade to be identified to obtain a blade entity model; dividing the cross-sectional profile of the blade entity model into regions, calculating corresponding airfoil shaping parameters according to the values ​​of the area of ​​each region, to obtain the airfoil shaping parameters; outputting the model of the NACA airfoil blade to be identified according to the airfoil shaping parameters; and performing a lofting operation on the cross-sectional profile according to the model to generate a NACA airfoil blade model. After scanning the NACA airfoil blade to be identified, the corresponding blade entity model is obtained, and the area method is used to divide the area of ​​the blade entity model, so as to facilitate the determination of the area of ​​the NACA airfoil blade to be identified, thereby facilitating the calculation of the corresponding airfoil shaping parameters according to the regional area distribution. Finally, the model can be determined according to the airfoil shaping parameters, thereby reducing the error caused by the blade profile. Moreover, only the area of ​​the blade needs to be calculated, thereby reducing the amount of data processing, effectively improving the accuracy of model identification, and reducing the data processing difficulty of reverse modeling of airfoil blades.

[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 part or all of the following steps.

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

[0055] In the present embodiment, the NACA airfoil blade to be identified is a blade that needs to be identified as the model 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 analysis and 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 point clouds of the blade solid surface, and pre-processes the data point clouds. With the help of three-dimensional solid modeling software and related technologies, the reconstruction of the blade solid model is realized, 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 each region to obtain the airfoil shaping parameters.

[0057] In this embodiment, the blade entity model is the three-dimensional object contour data of the NACA airfoil blade to be identified. The blade entity model is used as the surface contour of the NACA airfoil blade to be identified, which is convenient for determining the three-dimensional surface molding of the NACA airfoil blade to be identified. The area division of the cross-sectional contour of the blade entity model and the calculation of the corresponding airfoil shaping parameters according to the area values ​​of each area are to divide the cross-sectional contour of the blade entity model by the area method. Specifically, the cross-sectional contour of the blade entity model is divided into areas, and the corresponding airfoil shaping parameters are calculated by the divided areas, wherein the airfoil shaping parameters correspond to the contour features of the NACA airfoil blade to be identified, which is convenient for determining 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 entity model, and the blade entity model is the three-dimensional object contour data of the NACA airfoil blade to be identified. The blade entity model is used as the surface contour of the NACA airfoil blade to be identified, which is convenient for determining the three-dimensional surface molding of the NACA airfoil blade to be identified. The area division of the cross-sectional contour of the blade entity model and the calculation of the corresponding airfoil shaping parameters according to the area value of each area are to divide the cross-sectional contour of the blade entity model by the area method. Specifically, the cross-sectional contour of the blade entity model is divided by the area of ​​the airfoil contour, and the corresponding airfoil shaping parameters are calculated by the divided area, wherein the airfoil shaping parameters correspond to the contour features of the NACA airfoil blade to be identified, which is convenient for determining 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 curvature of the airfoil, and the position of the maximum curvature 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 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, the corresponding blade solid model is obtained. The area method is used to divide the blade solid model into regions, 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 according to the regional area distribution. Finally, the model can be determined according to the airfoil shaping parameters, thereby reducing the error caused by the blade contour. Moreover, only the area of ​​the blade needs to be calculated, thereby reducing 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.

[0062] In one embodiment, the said dividing the cross-sectional contour of the blade solid model into regions, and calculating the corresponding airfoil shaping parameters according to the area value of each region to obtain the airfoil shaping parameters, comprises: performing a cross-sectional operation on the blade solid model to obtain the airfoil contour. In this embodiment, the blade solid model is the three-dimensional object contour data of the NACA airfoil blade to be identified, and the blade solid model is used as the surface contour of the NACA airfoil blade to be identified, so as to facilitate the determination of the three-dimensional surface forming of the NACA airfoil blade to be identified. The said dividing the cross-sectional contour of the blade solid model into regions, and calculating the corresponding airfoil shaping parameters according to the area value of each region, is to divide the cross-sectional contour of the blade solid model into regions by the area method, specifically, to divide the cross-sectional contour of the blade solid model into areas, and to calculate the corresponding airfoil shaping parameters by the divided areas, wherein the airfoil shaping parameters correspond to the contour features of the NACA airfoil blade to be identified, so as to facilitate the determination of the airfoil characteristics of the NACA airfoil blade to be identified. In the process of dividing the cross-sectional contour into regions and calculating the corresponding airfoil shaping parameters according to the area values ​​of each region, the blade solid model is first subjected to a cross-sectional operation so as to slice the blade solid model to obtain the precise airfoil contour of the NACA airfoil blade to be identified. Specifically, the blade solid model is face aligned, i.e., 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 contour.

[0063] Further, the section operation of the blade solid model is performed 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 vertical line operation on the airfoil chord distribution line to obtain a chord vertical line passing through the upper and lower profile lines of the airfoil; and obtaining the four-image partition area of ​​the airfoil profile according to the airfoil chord distribution line and the chord vertical line. In this embodiment, the blade solid model is used as the surface profile of the NACA airfoil blade to be identified, which is convenient for determining the three-dimensional surface forming of the NACA airfoil blade to be identified. The area division of the cross-sectional profile of the blade solid model is to calculate the corresponding airfoil shaping parameters according to the value of the area of ​​each area, which is to divide the cross-sectional profile of the blade solid model by the area method. Specifically, the cross-sectional profile of the blade solid model is divided into the area of ​​the airfoil profile, and the corresponding airfoil shaping parameters are calculated by the divided area, wherein the airfoil shaping parameters correspond to the profile features of the NACA airfoil blade to be identified, which is convenient for determining 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, the blade entity model is first subjected to a cross-sectional operation so as to slice the blade entity model to obtain the precise airfoil profile of the NACA airfoil blade to be identified. Specifically, the blade entity model is face aligned, i.e., the plane at the top surface of the polyhedron is used as the reference plane for alignment. After alignment, the blade model is subjected to cross-sectional acquisition by an equidistant interception method to obtain the airfoil profile. The airfoil profile is the profile corresponding to the precise airfoil profile of the NACA airfoil blade to be identified. By operating the profile curvature of the airfoil profile, the chord of the airfoil blade is determined according to the profile curvature distribution of the NACA airfoil blade to be identified. Specifically, the point with the largest curvature at the rear end is found according to the curvature calculation, i.e., the tail of the airfoil, and it is used 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 with 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 rear 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 the chord distribution line, i.e., the chord of the airfoil blade, is determined, the vertical line of the chord distribution line of the airfoil is operated to make the corresponding chord vertical line based on the chord, and the chord vertical line passes through the upper and lower contour lines of the airfoil profile. For example, the midpoint of the chord distribution line of the airfoil is selected as the vertical point of the chord vertical line. In this way, after the chord and the chord vertical line are determined, the chord and the chord vertical line divide the airfoil profile into regions to form the four-image partition area of ​​the airfoil profile. Specifically, the four-image partition area of ​​the airfoil profile is the four divided areas formed on the airfoil profile, so as to facilitate the area calculation of each area on the airfoil profile.

[0065] Furthermore, the method further comprises: performing a merging and integrating 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 and integrating 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 area in 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 It is the area of ​​the lower right side of the four-image partition area of ​​the airfoil contour, 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 a chord length c. Therefore, in this coordinate system, the coordinate of the trailing edge B is (1, 0). c is the curvature value d about the change of 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, parameter m is the maximum curvature, parameter p is the position of the maximum curvature, and 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 line of the airfoil, y l is the coordinate y value of the lower profile line of the airfoil. 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 blade to be identified is calculated by 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 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 obtaining of 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 area in 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It 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 mid-camber line, which is a curve connecting the front and rear 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 mid-camber line. For details, see 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 expressed 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 is expressed as:

[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 area in 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 It is the area of ​​the lower right side of the four-image partition area of ​​the airfoil contour, and p is the position of the maximum curvature of the airfoil.

[0094] In this embodiment, according to formula (11), it can be obtained:

[0095]

[0096] Because of y c It is a piecewise function. Using either formula (14) or formula (15) alone can obtain the area formula for parameter p. However, the selection of the maximum curvature position p of the airfoil is segmented. Therefore, the two formulas are considered separately. 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 get an area formula for parameter p, which is convenient for obtaining the maximum curvature position p of the airfoil by integrating the four-image partition area 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, we only need to get 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 region, 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 "stretching" 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 known 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 can 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 edge of the curved surface 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 operation of the NACA airfoil blade reverse modeling method is as follows:

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

[0102] Step 2. Preprocess 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. See the attached figure for details. Figure 7 .

[0104] Step 4. Find the chord through curvature analysis and tangent points, draw the perpendicular midline of the chord, 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 Fig. 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 Fig.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 value into the airfoil area formula containing parameters t and m, we can get t = 0.099876, m = 0.020492, and rounding t and m to get t = 0.10, m = 0.02. Substituting the area value and m into the airfoil area formula containing parameter p, we can get p1 = 0.766288, p2 = 0.418982, p3 = 0.332248 + 0.018990i, p4 = 0.332248 - 0.018990i. Since p should be within the range and be a real number, p = p2 = 0.418982, and rounding p to get p = 0.4, that is, the model of 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 Fig.11 As shown. The lofting operation for the cross-sectional profile is to bring the model into the above airfoil formula so as to inversely calculate 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 deviations of other areas are mainly yellow, and there is no red with large deviations, which meets the tolerance requirements of the blade. Fig.12 shown.

[0110] In one embodiment, the NACA airfoil blade reverse modeling method involved in the present disclosure has differences in airfoil parameter calculation compared with 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 scheme draws the median arc line through the midpoint of the ten-equal points, such as Fig.13 As shown, the maximum distance point between the mid-arc and the chord is found through the tangent relationship, as Fig.14 As shown, finally passing through the tangency point, the software measured m=0.019515 and 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 inserts them into the area formula to obtain m=0.019552, p=0.385927.

[0113] It can be seen that the existing solution has Δm=0.000485 and Δp=0.017468, while the disclosed solution 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, and the operation is simpler.

[0115] In one of the embodiments, 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 determine the regional area of ​​the NACA airfoil blade to be identified, thereby facilitating the calculation of the corresponding airfoil shaping parameters according to the regional area distribution. Finally, the airfoil blade reverse generation module can determine the model according to the airfoil shaping parameters, thereby reducing the error caused by the blade contour. Moreover, only the area of ​​the blade needs to be calculated, thereby reducing the amount of data processing, effectively improving the accuracy of model recognition and reducing 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: Fig.15As shown. The computer device includes a processor, a memory and a network interface connected through a system bus. Among them, 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 through 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 Fig.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 those shown in the figure, or combine certain components, or have a different arrangement of components.

[0119] In one of the embodiments, 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 of the embodiments, the present application further provides a computer-readable storage medium having a computer program stored thereon, and when the computer program is executed by a processor, the steps in the above-mentioned method embodiments are implemented.

[0121] Those of ordinary skill in the art can understand that all or part of the processes in the above-mentioned embodiment methods can be completed 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 can include at least one of non-volatile and volatile memory. Non-volatile memory can include read-only memory (ROM), magnetic tape, floppy disk, flash memory or optical memory, etc. Volatile memory can 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-mentioned embodiments only express several implementation methods of the present disclosure, and the descriptions thereof are relatively specific and detailed, but they cannot be understood as limiting the scope of the invention patent. It should be pointed out that, for ordinary technicians in this field, several variations and improvements can be made without departing from the concept of the present disclosure, and these all belong to the protection scope of the present disclosure. Therefore, the protection scope of the patent of the present disclosure shall be subject to the attached claims.

Claims

1. A method for reverse modeling of a NACA airfoil blade, 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; A lofting operation is performed on the cross-sectional profile according to the model to generate a NACA airfoil blade model.

2. The NACA airfoil blade reverse modeling method according to claim 1, characterized in that: 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: A cross-section operation is performed on the blade solid model to obtain an airfoil profile.

3. The NACA airfoil blade reverse modeling method according to claim 2, characterized in that: The step of performing a cross-section operation on the blade solid model to obtain an airfoil profile further comprises: 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; The four-dimensional partition area of ​​the airfoil profile is obtained according to the airfoil chord distribution line and the chord perpendicular line.

4. The NACA airfoil blade reverse modeling method according to claim 3, characterized in that: The performing a vertical line operation on the airfoil chord distribution line comprises: The midpoint of the airfoil chord distribution line is selected as the vertical point of the chord perpendicular line.

5. 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; and then: The four-image partition area of ​​the airfoil profile is combined and integrated to obtain the maximum thickness of the airfoil. The calculation formula of the combined and integrated 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 area in 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 It is the area of ​​the lower right side of the four-image partition area of ​​the airfoil contour, and t is the maximum thickness of the airfoil.

6. 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; and then: The difference integral operation is performed on the four-image partition area of ​​the airfoil profile to obtain the maximum curvature of the airfoil. The calculation formula of the difference integral operation satisfies the following conditions: Among them, S 左上 is the area of ​​the upper left area in 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 It 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.

7. The NACA airfoil blade reverse modeling method according to claim 6, 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; and then: 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 satisfies the following conditions: Among them, S 左上 is the area of ​​the upper left area in 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 It is the area of ​​the lower right side of the four-image partition area of ​​the airfoil contour, and p is the position of the maximum curvature of the airfoil.

8. A NACA airfoil blade model identification device, characterized in that: include: An airfoil blade acquisition module, wherein the airfoil blade acquisition module is used to perform blade scanning on 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 the 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 lofting operations on the cross-sectional profile according to the model to generate a NACA airfoil blade model.

9. 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 7 are implemented.

10. 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 7 are implemented.

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