Aero-engine damaged blade data repairing and modeling method

Abstract

The invention provides an aero-engine damaged blade data restoration and modeling method, which comprises the following steps: processing finite element grid files of a prototype blade and a damaged blade, and extracting grid topology information and node information of the blade; comparing node information of the prototype blade and the damaged blade, and determining multiple types of damaged nodes in the damaged blade; estimating the correct node position of the damaged node through three-dimensional data interpolation, and repairing the damaged area; the intermediate state shape of the blade is reconstructed by setting a repair reduction coefficient; and carrying out secondary reconstruction on nodes of the front edge area and the tail edge area of the blade, arranging blade profile node coordinates, and outputting a blade surface blade profile geometric file supporting pneumatic simulation. According to the method, the automatic conversion process from the finite element simulation grid to the pneumatic simulation blade profile file can be efficiently achieved, the blade damage analysis and repair design efficiency is greatly improved, and subsequent pneumatic performance simulation is facilitated.

Description

Technical Field

[0001] The present application relates to the technical field of aero-engine blade design, and in particular to a method for repairing and modeling data of damaged aero-engine blades. Background Art

[0002] Currently, aircraft engines frequently ingest environmental debris, such as metal objects, sand, gravel, birds, and hail, during takeoff, landing, or airborne navigation. Foreign object damage (FOD) occurs when foreign objects enter an aircraft engine through the airflow and impact the fan blades. During the life of an aircraft, all types of aircraft are at high risk of experiencing FOD. Literature and statistics indicate that FOD incidents are a frequent occurrence on aircraft engine blades, and the number of engines requiring repairs due to FOD incidents accounts for a significant portion of the total number of repairs.

[0003] When an aircraft engine fan is damaged by a foreign object, it may cause the fan blades to deform, tear, fall off, or even break, which will directly affect the normal operation of the fan, resulting in a decrease in the efficiency of introducing and compressing air, affecting the overall performance of the engine. In addition, the impact of foreign objects may cause the vibration of the fan blades to intensify, triggering a series of chain reactions. For example, it may cause the stator and rotor to contact, increase the amount of air leakage, and thus affect the operating efficiency of the engine. The vibration may also cause damage to other engine components, further exacerbating the damage to the engine. Furthermore, the engine material that falls off after the impact flows with the airflow, further damaging other fan blades and other engine components, causing the engine to stall and increase the vibration of the high-pressure rotor. In severe cases, the engine may not be able to maintain normal takeoff thrust, resulting in a significant loss of thrust.

[0004] In related technologies, when repairing damage caused by foreign objects hitting fan blades, the finite element method is generally used to simulate the damaged state of the blades first, obtain the finite element mesh data of the blades after deformation, tearing, and falling off, and then conduct an aerodynamic assessment of the damaged blades.

[0005] However, the aforementioned related art solutions generally require manual extraction of data points from finite element mesh files when evaluating aerodynamic performance, and lack tools for automatically converting deformed and damaged blade finite element mesh files into aerodynamic simulation meshes. Furthermore, reverse engineering repair of chipped or torn finite element mesh files typically requires manual point addition and modeling, a time-consuming and labor-intensive process with low accuracy. Automated design methods and tools are also lacking for reverse engineering repair of chipped or torn finite element mesh files.

[0006] Therefore, how to efficiently and accurately realize the conversion of damaged aircraft engine fan blades from finite element mesh files to aerodynamic simulation mesh files has become an urgent problem that needs to be solved. Summary of the Invention

[0007] The present application aims to solve one of the technical problems in the related art at least to a certain extent.

[0008] To this end, the first purpose of this application is to propose a method for data repair and modeling of damaged blades of aircraft engines. This method can automatically detect and repair damaged blades after finite element calculation in a more efficient manner, and realize the automated conversion process from finite element simulation mesh to aerodynamic simulation blade profile file, greatly improving the efficiency of blade damage analysis and repair design, facilitating subsequent aerodynamic performance simulation in flow field numerical simulation software such as NUMECA, and realizing the repair and modeling of torn and chipped blade profile mesh files.

[0009] The second purpose of this application is to propose a data repair and modeling system for damaged aircraft engine blades.

[0010] A third object of the present application is to provide a non-transitory computer-readable storage medium.

[0011] To achieve the above objectives, the first aspect of the present application is to propose a method for repairing and modeling damaged aero-engine blade data, comprising the following steps:

[0012] The finite element mesh files of the prototype and damaged blades of the aircraft engine are processed to extract the mesh topology and node information of each blade;

[0013] comparing the node information of the prototype blade and the damaged blade to determine multiple types of damaged nodes in the damaged blade;

[0014] Based on the mapping relationship between the prototype blade nodes and the damaged blade nodes, the correct node position of the damaged node is estimated through three-dimensional data interpolation to repair the damaged area of the blade;

[0015] By setting a repair restoration coefficient, the deflection state of the damaged blade after finite element calculation is repaired to reconstruct the intermediate shape of the blade;

[0016] The nodes in the leading and trailing edge areas of the damaged blade after deflection repair are reconstructed, and the blade profile node coordinates obtained after multiple rounds of repair are sorted out to output a blade profile geometry model file that supports aerodynamic simulation.

[0017] Optionally, in one embodiment of the present application, the extracting of mesh topology information and node information of each blade includes: according to the storage format of the finite element mesh file, dividing the file data into hexahedral or tetrahedral data format, and extracting the volume mesh topology information and the surface mesh topology information; according to the volume mesh topology information and the surface mesh topology information, using the lower meridian to filter out the nodes on the surface profile of the blade part; identifying multiple key repair nodes on the blade surface, and arranging the nodes of each layer of the blade in order.

[0018] Optionally, in one embodiment of the present application, determining multiple types of damaged nodes in the damaged blade includes: arranging the node data of the prototype blade and the damaged blade in the same format and order to achieve a one-to-one correspondence between the nodes in the two blades; respectively calculating the Euclidean distance in three-dimensional space for each pair of nodes in the prototype blade and the damaged blade; and combining the Euclidean distance, radial blade height and pitch to determine the nodes that have suffered different types of damage.

[0019] Optionally, in one embodiment of the present application, the correct node position of the damaged node is estimated through three-dimensional data interpolation based on the mapping relationship between the prototype blade node and the damaged blade node, including: taking the undamaged nodes in the prototype blade as a sample point group for repair reference, and taking the damaged nodes in the damaged blade as an experimental point group to be repaired; based on the correspondence between the sample point group and the experimental point group and the parameters of the surrounding nodes of the damaged node, three-dimensional interpolation repair is performed on each damaged node in the experimental point group, wherein, in the process of the three-dimensional interpolation repair, a weighted average calculation method is used to interpolate the spatial coordinates.

[0020] Optionally, in one embodiment of the present application, the repairing of the deflection state of the damaged blade after finite element calculation includes: calculating the deflection vector of each node by comparing the node coordinates of the prototype blade and the deformed blade after finite element calculation; setting the repair restoration coefficient according to the recovery degree requirement, and adjusting the size of the deflection vector by the repair restoration coefficient.

[0021] Optionally, in one embodiment of the present application, the nodes of the leading edge area and the trailing edge area of the damaged blade after deflection repair are reconstructed secondary, including: setting the leading edge repair point and the trailing edge repair point, extracting the node coordinates of the leading edge area and the trailing edge area of each layer of the blade profile of the damaged blade; converting the node coordinates of the leading edge area and the trailing edge area located in the three-dimensional space to the XY plane of the Cartesian coordinate system to realize the local coordinate projection transformation; calling the non-uniform rational Bezier NURBS curve to reconstruct the transformed leading edge area and the trailing edge area.

[0022] Optionally, in one embodiment of the present application, after the blade profile node coordinates obtained after multiple rounds of repair are sorted out, it also includes: redistributing the sorted blade profile node coordinates along the radial direction; the blade surface profile geometry model that supports aerodynamic simulation includes: a blade profile geometry file in geomturbo format.

[0023] To achieve the above objectives, the second aspect of the present application further proposes a data repair and modeling system for damaged aircraft engine blades, comprising the following modules:

[0024] The extraction module is used to process the finite element mesh files of the prototype blades and damaged blades of the aircraft engine and extract the mesh topology information and node information of each blade;

[0025] a determination module, configured to compare the node information of the prototype blade and the damaged blade, and determine multiple types of damaged nodes in the damaged blade;

[0026] A first repair module is configured to estimate the correct node position of the damaged node through three-dimensional data interpolation based on the mapping relationship between the prototype blade node and the damaged blade node, so as to repair the damaged area of the blade;

[0027] a reconstruction module, configured to repair the deflection state of the damaged blade after finite element calculation by setting a repair restoration coefficient, so as to reconstruct the intermediate shape of the blade;

[0028] The second repair module is used to reconstruct the nodes in the leading edge area and trailing edge area of the damaged blade after deflection repair, organize the blade profile node coordinates obtained after multiple rounds of repair, and output the blade profile geometry model file of the blade surface that supports aerodynamic simulation.

[0029] In order to implement the above-mentioned embodiments, the third aspect of the present application further proposes a non-temporary computer-readable storage medium on which a computer program is stored. When the computer program is executed by a processor, the method for repairing and modeling damaged aircraft engine blade data in the first aspect mentioned above is implemented.

[0030] The technical solutions provided by the embodiments of this application offer at least the following beneficial effects: This application enables automated repair and performance evaluation of damaged aircraft engine fan blades, enabling automated conversion from finite element simulation meshes to aerodynamic simulation blade profile files. This conversion process is both efficient and accurate, significantly improving the efficiency of blade damage analysis and repair design. The program's technical framework incorporates a reverse repair function for blade profile mesh files for typical damage types, such as tears and chipping. This function automatically detects and repairs the mesh in damaged areas, while also allowing for customized settings for the repair restoration effect. This ensures that the repaired blade profile meets aerodynamic performance requirements while maintaining the integrity of the original design. This allows for visualization of the blade repair process, ensuring repeatability and standardization of the entire repair process, and effectively avoiding errors potentially caused by human intervention. This application can significantly shorten the design cycle for aircraft engine blades and significantly reduce time costs. Furthermore, this application significantly reduces the high cost and resulting accuracy issues associated with traditional blade damage analysis, which relies on scanning and sampling test pieces. Directly and accurately modeling and analyzing damaged blades through digital means not only improves the accuracy of the assessment but also greatly reduces costs, providing a more economical, efficient and reliable solution for the aerodynamic performance analysis and evaluation of damaged aircraft engine blades.

[0031] Additional aspects and advantages of the present invention will be set forth in part in the description which follows and, in part, will be obvious from the description which follows, or may be learned through practice of the present invention. BRIEF DESCRIPTION OF THE DRAWINGS

[0032] The above and / or additional aspects and advantages of the present application will become apparent and easily understood from the following description of the embodiments in conjunction with the accompanying drawings, in which:

[0033] Figure 1 This is a flow chart of a method for repairing and modeling damaged aero-engine blade data proposed in an embodiment of the present application;

[0034] Figure 2 A schematic diagram of a process for extracting and organizing blade grid topology data and node data proposed in an embodiment of the present application;

[0035] Figure 3 A schematic diagram of a blade damage form proposed in an embodiment of the present application;

[0036] Figure 4 A schematic diagram of a method for determining a damaged node proposed in an embodiment of the present application;

[0037] Figure 5 A schematic diagram of a specific restoration process based on three-dimensional spatial data interpolation proposed in an embodiment of the present application;

[0038] Figure 6 This is a schematic diagram of the repair status of damaged blades in various process states under different repair and restoration coefficients proposed in an embodiment of the present application;

[0039] Figure 7 This is a structural diagram of a data repair and modeling system for damaged aircraft engine blades proposed in an embodiment of the present application. DETAILED DESCRIPTION

[0040] The following describes embodiments of the present invention in detail, examples of which are shown in the accompanying drawings, wherein the same or similar reference numerals throughout represent the same or similar elements or elements having the same or similar functions. The embodiments described below with reference to the accompanying drawings are exemplary and are intended to be used to explain the present invention, and are not to be construed as limiting the present invention.

[0041] The following describes, with reference to the accompanying drawings, a method and system for repairing and modeling damaged aero-engine blade data proposed in an embodiment of the present application.

[0042] Figure 1 This is a flow chart of a method for repairing and modeling damaged aero-engine blade data proposed in an embodiment of the present application, as shown in FIG. Figure 1 As shown, the method includes the following steps:

[0043] Step S101 : Processing the finite element mesh files of the prototype blade and the damaged blade of the aircraft engine to extract mesh topology information and node information of each blade.

[0044] Specifically, the blade mesh topology and node data are extracted before and after the finite element calculation. This involves extracting the topological identification information and node information of the hexahedral or tetrahedral mesh of the engine blade before and after damage. This information extraction process includes identifying the information and organizing the blade nodes.

[0045] In one embodiment of the present application, the mesh topology information and node information of each blade are extracted, including: according to the storage format of the finite element mesh file, the file data is divided into a hexahedron or tetrahedron data format, and the volume mesh topology information and the surface mesh topology information are extracted; according to the volume mesh topology information and the surface mesh topology information, the nodes on the surface profile of the blade part are screened out using the lower meridian; multiple key repair nodes on the blade surface are identified, and the nodes of each layer of the blade are arranged in order.

[0046] Specifically, the information extraction process of this embodiment is as follows: Figure 2As shown in the figure, after importing the finite element mesh files of the prototype and damaged blades into the automated programming program, the data is automatically classified into hexahedral or tetrahedral data formats based on the finite element mesh file storage encoding format, and the mesh topology information and node space data are extracted. Among them, the common topology information includes surface mesh and volume mesh, corresponding to the two element structures of CQUAD4 and CHEXA, respectively.

[0047] After obtaining the topological information for the volume and surface meshes, the program uses the meridian flow path to segment the blade's airfoil nodes. Based on this topological information, internal nodes are excluded. The dataset is then partitioned using the lower meridian spatial coordinates to identify nodes on the blade's airfoil surface profile. The program then identifies and sorts several specific nodes on the blade's surface, representing key repair nodes, such as the top profile nodes, leading edge nodes, and trailing edge nodes. The program then searches layer by layer from the blade tip to the blade root, achieving an orderly arrangement of nodes at each layer.

[0048] Therefore, this step determines the node information of the leading and trailing edges of the blade and the top blade profile line based on the volume mesh and surface mesh topology information, and obtains the orderly arrangement of the blade profile lines of each layer.

[0049] Step S102 : comparing the node information of the prototype blade and the damaged blade to determine various types of damaged nodes in the damaged blade.

[0050] Specifically, the spatial positions of the data nodes of the intact blade and the damaged blade are compared to find the loss area of the blade. By calculating the Euclidean distance, the various types of damaged nodes in the damaged blade are determined.

[0051] Among them, various types of damage include deformation, tearing, dislocation, and falling pieces, which may occur to blades due to FOD events.

[0052] For example, if Figure 3 As shown in the figure, after finite element calculations, a damaged blade may exhibit significant geometric deviations from the prototype blade. These deviations may manifest as nodes that are crossed, misaligned, or protruding outward, causing these nodes to no longer lie on the normal blade surface. Therefore, it is necessary to compare the positions of the nodes on the prototype and damaged blades and use data calculations to determine which nodes have been damaged, such as those that have been destroyed or displaced.

[0053] In one embodiment of the present application, multiple types of damaged nodes in a damaged blade are determined, including: arranging the node data of the prototype blade and the damaged blade in the same format and order to achieve a one-to-one correspondence between the nodes in the two blades; calculating the Euclidean distance of each pair of nodes in the prototype blade and the damaged blade in three-dimensional space; and combining the Euclidean distance, radial blade height and pitch to determine the nodes that have suffered different types of damage.

[0054] Specifically, in order to accurately identify the problem node in this embodiment, the node data of the prototype blade and the deformed blade are compared, and the Euclidean distance between the corresponding nodes is mainly calculated to determine whether the node has been deformed or fallen off. Figure 4 The judgment principle shown in FIG. 5 and the specific steps for judging the damaged node are as follows.

[0055] The first step is to ensure the consistency of the data matrix. Arrange the node data of the prototype blade and the damaged blade in the same format and order to ensure a one-to-one correspondence between the topological relationships between the two. That is, each node in the prototype blade has a corresponding node position in the deformed blade. For example, Figure 4 The prototype leaf node 10 and the damaged leaf judgment node 20 are shown in FIG.

[0056] The second step is to calculate the Euclidean distance. For each pair of prototype leaf nodes and damaged leaf nodes, calculate their Euclidean distance in three-dimensional space. The Euclidean distance can be calculated using the following formula:

[0057]

[0058] Among them, (x orig ,y orig , z orig ) and (x damaged ,y damaged , z damaged ) are the three-dimensional coordinates of the nodes of the prototype blade and the damaged blade, respectively.

[0059] The third step is to preliminarily identify damaged nodes. By calculating the Euclidean distance, if the calculated Euclidean distance exceeds a certain threshold, the node is considered to have undergone significant displacement or damage. Based on this, and considering the physical characteristics of the blade after deformation, the radial height and pitch of the blade can be used to further restrict and verify the rationality of the node.

[0060] The fourth step is to make a final judgment based on the radial blade height and pitch. Based on the geometric characteristics of the blade and the radial height and pitch of the nodes, the position of each node within its corresponding area can be determined. By calculating the pitch information of each node of the prototype blade at different radial heights and comparing it with the node information of the deformed blade, we can further identify nodes that are missing or misaligned.

[0061] Therefore, this step identifies nodes in the broken blade that are damaged, dislocated, or detached by comparing and analyzing the prototype blade data and the broken blade data.

[0062] Step S103 , based on the mapping relationship between the prototype blade node and the damaged blade node, the correct node position of the damaged node is estimated by three-dimensional data interpolation to repair the damaged area of the blade.

[0063] Specifically, based on the mapping relationship between the prototype blade nodes and the damaged blade nodes, the correct node position of the damaged area is inferred through three-dimensional data interpolation to achieve the repair of the damaged area of the blade. The main purpose of the repair process of this application is to use the spatial information of the undamaged nodes to estimate the position of the block nodes. Specifically, it is to establish a corresponding relationship between the prototype blade data points and the damaged blade data points, and repair the block nodes of the damaged blade according to the relationship.

[0064] In one embodiment of the present application, based on the mapping relationship between the prototype blade node and the damaged blade node, the correct node position of the damaged node is estimated by three-dimensional data interpolation, including: using the undamaged nodes in the prototype blade as a sample point group for repair reference, and using the damaged nodes in the damaged blade as an experimental point group to be repaired; based on the correspondence between the sample point group and the experimental point group and the parameters of the surrounding nodes of the damaged node, three-dimensional interpolation repair is performed on each damaged node in the experimental point group, wherein a weighted average calculation method is used to interpolate the spatial coordinates during the three-dimensional interpolation repair process.

[0065] For example, taking the repair of lost nodes as an example, the repair of the damaged area in this embodiment includes the following steps: The first step is to establish a node mapping relationship. For the prototype blade and the damaged blade, there is a one-to-one correspondence between the nodes in the spatial area. The nodes that have not lost blocks are regarded as the "sample point group", that is, the known node set; and the lost nodes are regarded as the "experimental point group", that is, the node set that needs to be repaired. In other words, the nodes in the "sample point group" are regarded as prototype data nodes, and are used as repair reference data, and the lost nodes are regarded as target nodes to be repaired.

[0066] The second step is to perform three-dimensional interpolation mapping and use interpolation methods to repair the broken nodes in the "experimental point group". Common interpolation methods include linear interpolation and spline interpolation. By selecting an appropriate interpolation algorithm, the correct position of the broken node can be estimated based on the spatial position and attributes (such as node coordinates, topological position, etc.) of the known nodes (i.e., the undamaged nodes in the prototype blade). During the interpolation process, the surrounding nodes of the broken node can be considered as the basis for interpolation, and the spatial coordinates can be interpolated by weighted averaging and other methods.

[0067] The third step is to calculate the interpolation. For each missing node, the exact position in 3D space is calculated by interpolation based on the coordinates of the surrounding known nodes. This process is similar to using a weighted average method to infer the coordinates of missing points.

[0068] for example, Figure 5 The repair process of three-dimensional spatial data interpolation is shown. Figure 5 As you can see, the missing nodes (such as the red squares in the right image) are restored to their correct positions through interpolation. This method uses information from surrounding known nodes to infer the missing points. The repaired nodes restore the integrity of the blade surface.

[0069] Therefore, this step is based on the mapping relationship between the prototype blade nodes and the damaged blade nodes. Through three-dimensional data interpolation, the spatial information of the undamaged nodes is used to estimate the position of the block nodes, thereby realizing the repair of the damaged area of the blade.

[0070] Step S104 , repairing the deflection state of the damaged blade after finite element calculation by setting a repair restoration coefficient, so as to reconstruct the intermediate shape of the blade.

[0071] Specifically, this step aims to restore the blade's proper shape and provide preliminary geometric data required for aerodynamic simulation. By setting the restoration coefficient, the blade's distortion and deflection after finite element strength calculations are adjusted. Based on this, an interpolation or restoration algorithm is used to reconstruct the intermediate shape of the blade.

[0072] It should be noted that in finite element analysis, after the blade is calculated for strength, its shape will usually be distorted or deflected. This deflection is caused by the blade being subjected to external forces (such as aerodynamic forces, centrifugal forces, and temperature changes, etc.), reflecting the actual deformation of the blade. However, in aerodynamic simulation, excessive deflection will lead to inaccurate flow field predictions, because the shape of the blade and the direction of the airflow may have seriously deviated from the original design. Therefore, when performing aerodynamic simulation, it is often necessary to restore the approximate initial state of the blade, especially the intermediate shape in the strength calculation. This intermediate blade shape can avoid airflow deviation, thereby improving the accuracy and stability of the simulation calculation. In order to achieve this goal, the method provided in the embodiment of the present application introduces a repair restoration coefficient, and the intermediate shape of the blade is constructed by correcting the deflection vector between the prototype blade and the deformed blade.

[0073] In one embodiment of the present application, the deflection state of the damaged blade after finite element calculation is repaired, including: calculating the deflection vector of each node by comparing the node coordinates of the prototype blade and the deformed blade after finite element calculation; setting a repair restoration coefficient according to the recovery degree requirement, and adjusting the size of the deflection vector by the repair restoration coefficient.

[0074] Specifically, in this embodiment, the setting of the restoration coefficient is generally based on the following two aspects.

[0075] First, the comparison between the prototype blade and the deformed blade.

[0076] By comparing the node coordinates of the original blade and the twisted blade, the deflection vector of each node can be calculated. The deflection vector is the spatial difference between the deformed blade node and the original blade node, reflecting the deformation of the blade under the action of external forces. The calculation of the deflection vector can be expressed as follows:

[0077]

[0078] in, is the deflection vector, is the coordinate of the node after deformation, are the coordinates of the prototype leaf node.

[0079] Secondly, an appropriate restoration coefficient is introduced according to the degree of restoration required.

[0080] The restoration coefficient is a factor that adjusts the magnitude of the deflection vector. By setting this coefficient, the degree of restoration during the repair process can be controlled. Specifically, the restoration coefficient α controls the proportion of the shape of the repaired blade compared to the fully deformed blade calculated using finite element analysis. Typically, the restoration coefficient ranges from 0 to 1.

[0081] When α = 1, it means that the deformed blade is completely preserved, that is, no repair is performed;

[0082] When α = 0, it means that the shape of the original blade is completely restored;

[0083] When 0<α<1, it indicates partial recovery, i.e., the deflection is partially repaired.

[0084] For example, Figure 6 The program demonstrates the repair of damaged blades at various stages under different restoration coefficients. By adjusting the restoration coefficient, the program can appropriately restore the blade's deformation, resulting in an "intermediate" blade shape. This shape approximates the state at a certain moment in the finite element calculation process and is generally more suitable for aerodynamic simulation than the final deformed state. This intermediate blade is typically used for preliminary calculations in aerodynamic simulations, providing more reasonable blade geometry data for subsequent aerodynamic optimization and performance evaluation.

[0085] Step S105 , reconstructing the nodes of the leading edge region and the trailing edge region of the damaged blade after deflection repair, sorting out the blade profile node coordinates obtained after multiple rounds of repair, and outputting a blade profile geometry model file supporting aerodynamic simulation.

[0086] Specifically, the repair points of the leading edge and the trailing edge are set to make a smooth transition between the leading edge area and the trailing edge area. By performing a secondary reconstruction of the nodes of the leading edge and the trailing edge areas, the nodes of the leading edge and the trailing edge areas are ensured to be accurately aligned, avoiding the overlap and unreasonable distribution of the lines of the leading edge and the trailing edge areas.

[0087] It should be noted that when a fan blade is hit or damaged by a foreign object, especially in the leading edge area of the blade, blade collapse or falling pieces often occur. This situation not only causes some blade nodes to be misaligned or crossed, but may even be impacted to the outside of the blade, seriously affecting the geometric shape and aerodynamic performance of the blade. Although the above embodiment uses methods such as three-dimensional spatial data mapping to repair damaged blades, the node sorting and spatial distribution in the leading edge area of the blade may still be chaotic, which brings difficulties to the subsequent aerodynamic mesh division and calculation. Especially when the leading edge points on the suction side and the pressure side are improperly sorted and the profile lines overlap with each other, the repaired aerodynamic mesh may encounter erroneous or unstable calculation results. Therefore, for these problems, the present application also takes special repair measures, especially the repair of the leading edge area. This repair process requires not only the correction of the damaged nodes, but also the distribution of data points in the leading and trailing edge areas to be more accurate and smooth, thereby providing high-quality geometric input data for aerodynamic simulation.

[0088] In one embodiment of the present application, the nodes of the leading edge area and the trailing edge area of the damaged blade after deflection repair are reconstructed secondary, including: setting the leading edge repair point and the trailing edge repair point, extracting the node coordinates of the leading edge area and the trailing edge area of each layer of the blade profile of the damaged blade; converting the node coordinates of the leading edge area and the trailing edge area located in the three-dimensional space to the XY plane of the Cartesian coordinate system to realize the local coordinate projection transformation; calling the non-uniform rational Bezier NURBS curve to reconstruct the transformed leading edge area and the trailing edge area.

[0089] Specifically, in this embodiment, setting the repair leading edge point refers to the first retained node position on the suction side and the pressure side. Assuming that the repair leading edge point is selected as 3, the program will start from the leading edge, repair the blade profiles of the first two node areas on the suction side and the pressure side, and repair and reconstruct these areas. At the same time, the trailing edge area will also be reconstructed to ensure that the data point distribution accuracy of the leading and trailing edge areas is the same. By setting the repair point, it is possible to ensure that the nodes in the leading edge area are accurately aligned, and to avoid overlapping and unreasonable distribution of the leading edge area profile lines. This operation can ensure the geometric consistency of the leading and trailing edge areas, and provide more reasonable blade geometry input for aerodynamic simulation. Then extract the node coordinates of the leading and trailing edge areas of the blade profile lines of each layer of the damaged blade

[0090] Since the leading edge area of the blade is usually in three-dimensional space, and the cross-section of the blade profile is often in the form of a spatial plane, the processing and reconstruction of data points becomes complicated. In order to efficiently and accurately reconstruct these areas, the embodiment of the present application adopts a coordinate projection transformation method to project the data points in the three-dimensional space (for example, the extracted node coordinates of the leading edge area and the trailing edge area of each layer of the blade profile) onto a two-dimensional plane for processing. The coordinate projection transformation is to transform the data points in the three-dimensional space into the XY plane of the Cartesian coordinate system, which simplifies the data reconstruction process. The specific operation is to use the coefficient matrix of the spatial coordinate system to map the original three-dimensional coordinate points to the XY plane for curve fitting and reconstruction in the plane. Through this coordinate projection method, the data points in the leading edge and trailing edge areas of the blade can be processed more easily in a two-dimensional plane, reducing the complexity of the three-dimensional data reconstruction process.

[0091] Then, to ensure that the repaired blade leading and trailing edge area has a smooth and natural curve shape, this embodiment also uses a non-uniform rational Bezier (NURBS) curve to fit and control this area. The NURBS curve is an extension of the Bezier curve. Its biggest feature is that each control point can be attached with a weight, thereby affecting the shape of the curve. By adjusting the weight of the control point, this embodiment can accurately control the shape of the curve and ensure the smoothness of the leading and trailing edge area data. The specific mathematical description can be expressed by the following formula:

[0092]

[0093] Among them, p i is the control point position vector, B i,n (t) is the Bernstein basis function.

[0094] It can be understood that, compared with the Bezier curve, the characteristic of the NURB curve is that each control point has a weight. The NURB curve can control the curve shape by the position of the control point, and can also control the curve shape by changing the weight of the control point: the larger the weight of a control point, the closer the curve is to this control point; when the weight of a control point is 0, the control point has no control effect on the curve; when the weights of all control points are 1, the NURB curve degenerates into a Bezier curve.

[0095] Therefore, this step performs secondary reconstruction on the nodes in the leading / trailing edge areas prone to collapse or block loss after finite element calculation to ensure the rationality and smoothness of the point distribution.

[0096] Furthermore, based on the repair results, the blade profile node coordinates are organized and a profile geometry file corresponding to the engine blade is output. Specifically, this step also organizes the blade profile node coordinates obtained after multiple rounds of repair, including three-dimensional data interpolation repair, intermediate shape reconstruction, and secondary reconstruction, as described in the above embodiment. This outputs a profile geometry file of the repaired blade surface in various formats. This file may include a repaired blade profile geometry model.

[0097] In one embodiment of the present application, after sorting out the blade profile node coordinates obtained after multiple rounds of repair, it also includes: redistributing the sorted blade profile node coordinates along the radial direction; the output blade surface profile geometry model that supports aerodynamic simulation includes: a blade profile geometry file in geomturbo format.

[0098] Specifically, the repaired blade profile node coordinates are sorted and redistributed radially, and the blade geometry file in geomturbo format is named and output to facilitate subsequent blade modeling or blade performance simulation.

[0099] Therefore, this application aims to solve the problem of converting damaged fan blades of aircraft engines from finite element mesh files to aerodynamic simulation mesh files, and formulates effective control and repair strategies to realize the repair and modeling functions of torn and chipped blade mesh files, and constructs a reverse modeling program that meets the aerodynamic simulation requirements of damaged fan blades.

[0100] In summary, the damaged aeroengine blade data repair and modeling method of the present invention enables automated repair and performance evaluation of damaged aeroengine fan blades. It enables the automated conversion of finite element simulation meshes into aerodynamic simulation blade profile files. This conversion process is both efficient and accurate, significantly improving the efficiency of blade damage analysis and repair design. The program's technical framework incorporates a reverse repair function for blade profile mesh files for typical damage types, such as tears and chipping. This function automatically detects and repairs the mesh in damaged areas, while also allowing for customized settings for the repair restoration effect. This ensures that the repaired blade profile meets aerodynamic performance requirements while maintaining the integrity of the original design. This allows for visualization of the blade repair process, ensuring repeatability and standardization throughout the repair process, and effectively avoiding errors potentially introduced by human intervention. This method can significantly shorten the design cycle for aeroengine blades and significantly reduce time costs. Furthermore, it significantly reduces the high cost and resulting accuracy issues associated with traditional blade damage analysis, which relies on scanning and sampling test pieces. Directly and accurately modeling and analyzing damaged blades through digital means not only improves the accuracy of the assessment but also greatly reduces costs, providing a more economical, efficient and reliable solution for the aerodynamic performance analysis and evaluation of damaged aircraft engine blades.

[0101] In order to implement the above embodiment, the present application also proposes a data repair and modeling system for damaged blades of an aero-engine. Figure 7 This is a structural diagram of a data repair and modeling system for damaged blades of an aircraft engine proposed in an embodiment of the present application, as shown in FIG. Figure 7 As shown, the system includes:

[0102] The extraction module 100 is used to process the finite element mesh files of the prototype blade and the damaged blade of the aircraft engine, and extract the mesh topology information and node information of each blade.

[0103] The determination module 200 is configured to compare the node information of the prototype blade and the damaged blade, and determine various types of damaged nodes in the damaged blade.

[0104] The first repair module 300 is used to estimate the correct node position of the damaged node through three-dimensional data interpolation based on the mapping relationship between the prototype blade node and the damaged blade node, so as to repair the damaged area of the blade.

[0105] The reconstruction module 400 is used to repair the deflection state of the damaged blade after finite element calculation by setting a repair restoration coefficient, so as to reconstruct the intermediate shape of the blade.

[0106] The second repair module 500 is used to reconstruct the nodes in the leading edge area and the trailing edge area of the damaged blade after deflection repair, organize the blade profile node coordinates obtained after multiple rounds of repair, and output the blade profile geometry model file of the blade surface that supports aerodynamic simulation.

[0107] Optionally, in one embodiment of the present application, the extraction module 100 is specifically used to: divide the file data into hexahedral or tetrahedral data formats according to the storage format of the finite element mesh file, and extract the volume mesh topology information and the surface mesh topology information; based on the volume mesh topology information and the surface mesh topology information, use the lower meridian to filter out the nodes on the surface profile of the blade part; identify multiple key repair nodes on the blade surface, and arrange the nodes of each layer of the blade in order.

[0108] Optionally, in one embodiment of the present application, the determination module 200 is specifically used to: arrange the node data of the prototype blade and the damaged blade in the same format and order to achieve a one-to-one correspondence between the nodes in the two blades; calculate the Euclidean distance of each pair of nodes in the prototype blade and the damaged blade in three-dimensional space; and combine the Euclidean distance, radial blade height and pitch to determine the nodes that have suffered different types of damage.

[0109] Optionally, in one embodiment of the present application, the first repair module 300 is specifically used to: use the undamaged nodes in the prototype blade as a sample point group for repair reference, and use the damaged nodes in the damaged blade as an experimental point group to be repaired; based on the correspondence between the sample point group and the experimental point group and the parameters of the surrounding nodes of the damaged node, perform three-dimensional interpolation repair on each damaged node in the experimental point group, wherein a weighted average calculation method is used to interpolate the spatial coordinates during the three-dimensional interpolation repair process.

[0110] Optionally, in one embodiment of the present application, the reconstruction module 400 is specifically used to: calculate the deflection vector of each node by comparing the node coordinates of the prototype blade and the deformed blade after finite element calculation; set the repair restoration coefficient according to the recovery degree requirement, and adjust the size of the deflection vector by the repair restoration coefficient.

[0111] Optionally, in one embodiment of the present application, the second repair module 500 is specifically used to: set the leading edge repair point and the trailing edge repair point, extract the node coordinates of the leading edge area and the trailing edge area of each layer of the blade profile of the damaged blade; convert the node coordinates of the leading edge area and the trailing edge area located in the three-dimensional space to the XY plane of the Cartesian coordinate system to realize the local coordinate projection transformation; call the non-uniform rational Bezier NURBS curve to reconstruct the transformed leading edge area and the trailing edge area.

[0112] Optionally, in one embodiment of the present application, the second repair module 500 is further configured to: redistribute the sorted coordinates of each blade profile node in the radial direction.

[0113] It should be noted that the aforementioned explanation of the embodiment of the method for repairing and modeling damaged aero-engine blade data is also applicable to the system of this embodiment and will not be repeated here.

[0114] In summary, the damaged aero-engine blade data repair and modeling system of the present embodiment significantly shortens the aero-engine blade design cycle and significantly reduces time costs. Furthermore, the system directly and accurately models and analyzes damaged blades through digital means, improving assessment accuracy while significantly reducing costs.

[0115] In order to implement the above-mentioned embodiments, the present application also proposes a non-temporary computer-readable storage medium on which a computer program is stored. When the computer program is executed by a processor, it implements the method for repairing and modeling damaged blade data of an aircraft engine as described in any one of the above-mentioned first aspect embodiments.

[0116] In the description of this specification, the description with reference to the terms "one embodiment", "some embodiments", "example", "specific example", or "some examples" means that the specific features, structures, materials or characteristics described in conjunction with the embodiment or example are included in at least one embodiment or example of the present application. In this specification, the schematic representations of the above terms do not necessarily refer to the same embodiment or example. Moreover, the specific features, structures, materials or characteristics described can be combined in any one or more embodiments or examples in a suitable manner. In addition, those skilled in the art can combine and combine different embodiments or examples described in this specification and features of different embodiments or examples without contradiction.

[0117] Furthermore, the terms "first" and "second" are used for descriptive purposes only and should not be construed as indicating or implying relative importance or implicitly specifying the number of the technical features being referred to. Thus, a feature defined as "first" or "second" may explicitly or implicitly include at least one of such features. Throughout the description of this application, "plurality" means at least two, for example, two, three, etc., unless otherwise specifically defined.

[0118] Any process or method description in a flowchart or otherwise described herein may be understood to represent a module, segment or portion of code comprising one or more executable instructions for implementing the steps of a custom logical function or process, and the scope of the preferred embodiments of the present application includes alternative implementations in which functions may be performed out of the order shown or discussed, including performing functions in a substantially simultaneous manner or in the reverse order depending on the functions involved, which should be understood by those skilled in the art to which the embodiments of the present application belong.

[0119] The logic and / or steps represented in the flowcharts or otherwise described herein, for example, can be considered as a sequenced list of executable instructions for implementing the logical functions, and can be embodied in any computer-readable medium for use by, or in conjunction with, an instruction execution system, apparatus, or device (e.g., a computer-based system, a system including a processor, or other system that can fetch and execute instructions from an instruction execution system, apparatus, or device). For purposes of this specification, a "computer-readable medium" can be any device that can contain, store, communicate, propagate, or transport a program for use by, or in conjunction with, an instruction execution system, apparatus, or device. More specific examples (a non-exhaustive list) of computer-readable media include the following: an electrical connection with one or more wires (electronic devices), a portable computer disk cartridge (magnetic device), random access memory (RAM), read-only memory (ROM), erasable and programmable read-only memory (EPROM or flash memory), fiber optic devices, and a portable compact disc read-only memory (CDROM). Furthermore, the computer-readable medium may even be paper or other suitable medium on which the program is printed, since the program may be obtained electronically, for example, by optically scanning the paper or other medium and then editing, interpreting or processing it in another suitable manner if necessary, and then storing it in a computer memory.

[0120] It should be understood that various parts of the present application can be implemented using hardware, software, firmware, or a combination thereof. In the above embodiments, multiple steps or methods can be implemented using software or firmware stored in a memory and executed by a suitable instruction execution system. For example, if implemented using hardware, as in another embodiment, any one of the following technologies known in the art or a combination thereof can be used to implement: a discrete logic circuit having a logic gate circuit for implementing a logic function on a data signal, an application-specific integrated circuit having a suitable combination of logic gate circuits, a programmable gate array (PGA), a field programmable gate array (FPGA), etc.

[0121] Those skilled in the art will understand that all or part of the steps in the method of the above embodiment can be completed by instructing related hardware through a program, and the program can be stored in a computer-readable storage medium. When the program is executed, it includes one or a combination of the steps of the method embodiment.

[0122] In addition, the functional units in the various embodiments of the present application may be integrated into a processing module, or each unit may exist physically separately, or two or more units may be integrated into a module. The above-mentioned integrated module may be implemented in the form of hardware or in the form of a software functional module. If the integrated module is implemented in the form of a software functional module and sold or used as an independent product, it may also be stored in a computer-readable storage medium.

[0123] The storage medium mentioned above may be a read-only memory, a magnetic disk, or an optical disk, etc. Although the embodiments of the present application have been shown and described above, it is understood that the above embodiments are exemplary and should not be construed as limiting the present application. Persons skilled in the art may make changes, modifications, substitutions, and variations to the above embodiments within the scope of the present application.

Claims

1. A method for repairing and modeling damaged aero-engine blade data, characterized in that: The following steps are involved: The finite element mesh files of the prototype and damaged blades of the aircraft engine are processed to extract the mesh topology and node information of each blade; comparing the node information of the prototype blade and the damaged blade to determine multiple types of damaged nodes in the damaged blade; Based on the mapping relationship between the prototype blade nodes and the damaged blade nodes, the correct node position of the damaged node is estimated through three-dimensional data interpolation to repair the damaged area of the blade; By setting a repair restoration coefficient, the deflection state of the damaged blade after finite element calculation is repaired to reconstruct the intermediate shape of the blade; The nodes in the leading and trailing edge areas of the damaged blade after deflection repair are reconstructed, and the blade profile node coordinates obtained after multiple rounds of repair are sorted out to output a blade profile geometry model file that supports aerodynamic simulation.

2. The method according to claim 1, characterized in that The extraction of grid topology information and node information of each blade includes: According to the storage format of the finite element mesh file, the file data is divided into a hexahedron or a tetrahedron data format, and the volume mesh topology information and the surface mesh topology information are extracted; According to the volume mesh topology information and the surface mesh topology information, nodes on the surface profile of the blade part are screened out using the lower meridian; Identify multiple key repair nodes on the blade surface and arrange the nodes of each layer of the blade in sequence.

3. The method according to claim 1, characterized in that The determining of multiple types of damaged nodes in the damaged blade includes: Arranging the node data of the prototype blade and the damaged blade in the same format and order to achieve a one-to-one correspondence between the nodes in the two blades; respectively calculating the Euclidean distance between each pair of nodes in the prototype blade and the damaged blade in the three-dimensional space; The nodes with different types of damage are determined by combining the Euclidean distance, radial leaf height and pitch.

4. The method according to claim 3, characterized in that The estimating the correct node position of the damaged node by three-dimensional data interpolation based on the mapping relationship between the prototype blade node and the damaged blade node includes: Using the undamaged nodes in the prototype blade as a sample point group for repair reference, and using the damaged nodes in the damaged blade as an experimental point group to be repaired; Based on the correspondence between the sample point group and the experimental point group and the parameters of the surrounding nodes of the damaged node, three-dimensional interpolation repair is performed on each damaged node in the experimental point group, wherein a weighted average calculation method is used to interpolate the spatial coordinates during the three-dimensional interpolation repair process.

5. The method according to claim 1, wherein The repairing of the deflection state of the damaged blade after finite element calculation includes: By comparing the node coordinates of the prototype blade and the deformed blade after finite element calculation, the deflection vector of each node is calculated; The repair and restoration coefficient is set according to the restoration degree requirement, and the size of the deflection vector is adjusted by the repair and restoration coefficient.

6. The method according to claim 1, wherein The secondary reconstruction of the nodes in the leading edge region and the trailing edge region of the damaged blade after deflection repair includes: Setting the leading edge repair point and the trailing edge repair point, and extracting the node coordinates of the leading edge area and the trailing edge area of each layer of the blade profile of the damaged blade; Converting the node coordinates of the leading edge region and the trailing edge region in the three-dimensional space to the XY plane of the Cartesian coordinate system to achieve local coordinate projection transformation; The non-uniform rational Bezier NURBS curve is called to reconstruct the transformed leading edge area and trailing edge area.

7. The method according to claim 1, characterized in that After arranging the blade profile node coordinates obtained after multiple rounds of repair, the method further includes: Redistribute the sorted coordinates of each blade profile node along the radial direction; The blade surface profile geometry model supporting aerodynamic simulation includes: a blade profile geometry file in the geomturbo format.

8. A data repair and modeling system for damaged aircraft engine blades, characterized in that: Includes the following modules: The extraction module is used to process the finite element mesh files of the prototype blades and damaged blades of the aircraft engine and extract the mesh topology information and node information of each blade; a determination module, configured to compare the node information of the prototype blade and the damaged blade, and determine multiple types of damaged nodes in the damaged blade; A first repair module is configured to estimate the correct node position of the damaged node through three-dimensional data interpolation based on the mapping relationship between the prototype blade node and the damaged blade node, so as to repair the damaged area of the blade; a reconstruction module, configured to repair the deflection state of the damaged blade after finite element calculation by setting a repair restoration coefficient, so as to reconstruct the intermediate shape of the blade; The second repair module is used to reconstruct the nodes in the leading edge area and trailing edge area of the damaged blade after deflection repair, organize the blade profile node coordinates obtained after multiple rounds of repair, and output the blade profile geometry model file of the blade surface that supports aerodynamic simulation.

9. The system according to claim 8, characterized in that The extraction module is specifically used to: According to the storage format of the finite element mesh file, the file data is divided into a hexahedron or a tetrahedron data format, and the volume mesh topology information and the surface mesh topology information are extracted; According to the volume mesh topology information and the surface mesh topology information, nodes on the surface profile of the blade part are screened out using the lower meridian; Identify multiple key repair nodes on the blade surface and arrange the nodes of each layer of the blade in sequence.

10. A non-transitory computer-readable storage medium having a computer program stored thereon, characterized in that: When the computer program is executed by a processor, the method for repairing and modeling damaged aircraft engine blade data according to any one of claims 1 to 7 is implemented.

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