An aircraft airfoil pressure data correction method, device, equipment and medium

Abstract

The application discloses an aircraft airfoil pressure data correction method and device, equipment and medium, relates to the technical field of pressure distribution load correction, and comprises the following steps: determining a target position where a minimum value in each target chordwise coordinate is located, rearranging airfoil pressure distribution calculation data based on the target position, and decomposing the rearranged data to obtain upper-surface pressure distribution calculation data and lower-surface pressure distribution calculation data; obtaining a target pressure coefficient calculation value of a test point position of a target airfoil, determining a pressure coefficient difference between pressure distribution test data of the target airfoil and the target pressure coefficient calculation value; correcting the airfoil pressure distribution calculation data to obtain first corrected calculation data; performing a local smoothing operation on second corrected calculation data obtained by correcting the first corrected calculation data, and obtaining airfoil corrected data of the target airfoil; and obtaining corrected airfoil pressure distribution data according to the airfoil corrected data of the target airfoil. The data correction accuracy is improved.

Description

Technical Field

[0001] This invention relates to the field of pressure distribution load correction technology, and in particular to a method, apparatus, equipment and medium for correcting airfoil pressure data of aircraft. Background Technology

[0002] Aeroelastic coupling calculations for aircraft require the aerodynamics to transfer loads to the structure, and the structure to transfer responses to the aerodynamics. This involves iterative calculations between the two physical fields, and the accuracy of the aerodynamic loads directly impacts the prediction accuracy of the coupling. Currently, aerodynamic loads are primarily obtained through surface pressure integration predicted by Computational Fluid Dynamics (CFD). However, due to factors such as model simplification and numerical errors, the calculated pressure distribution data still deviates somewhat from the actual values. In practical applications, aircraft designers also place greater emphasis on pressure distribution data obtained from wind tunnel tests. However, due to limitations in physical measurement conditions, pressure test data suffers from insufficient resolution, making it difficult to obtain the refined aerodynamic load data required for aeroelastic coupling analysis through integration.

[0003] Calculated pressure distribution data offers advantages such as high resolution and high integration accuracy, but its accuracy is subject to certain deviations. Pressure test data, on the other hand, provides a true simulation of the physical process and boasts high accuracy, but its resolution and integration accuracy are lower. Therefore, combining the advantages of both types of data to obtain accurate aerodynamic loads is an important approach. However, the strong discontinuities in pressure distribution and the limited amount of test data pose significant challenges to data correction methods. Airfoil pressure distribution data corrected using a single interpolation method suffers from local jumps or insufficient accuracy, making it difficult to meet the data accuracy requirements for engineering applications. Summary of the Invention

[0004] In view of this, the purpose of this invention is to provide a method, apparatus, device, and medium for correcting airfoil pressure data, which can compensate for the deficiencies of pressure test data and calculated data, maintaining both the high accuracy of wind tunnel test data and the high resolution characteristics of calculated data. Furthermore, it can accurately correct complex flow regions such as shock waves and boundary layer separation, solving the problem of discrepancies between pressure test data and calculated data under high temperature and high pressure conditions. The specific solution is as follows:

[0005] In a first aspect, this application discloses a method for correcting airfoil pressure data of an aircraft, including:

[0006] The acquired airfoil pressure distribution test data of the aircraft is sorted to obtain sorted data. The target position of the minimum value in each target chord coordinate is determined. Based on the target position, the airfoil pressure distribution calculation data is rearranged to obtain rearranged data. The target chord coordinates are obtained by normalizing the airfoil pressure distribution calculation data of the aircraft.

[0007] The rearranged data is decomposed to obtain the pressure distribution calculation data of the upper wing surface and the pressure distribution calculation data of the lower wing surface;

[0008] The calculated value of the target pressure coefficient at the test point location of the target wing surface is obtained based on the pressure distribution calculation data of the target wing surface, and the pressure coefficient difference between the pressure distribution test data of the target wing surface and the calculated value of the target pressure coefficient in the sorted data is determined; the target wing surface is the upper wing surface or the lower wing surface;

[0009] The pressure coefficient difference is extended to obtain an extended difference. The extended difference is then smoothed to decompose it into a smooth distribution part and a non-smooth distribution part. The airfoil pressure distribution calculation data is then corrected based on the smooth distribution part to obtain the first corrected calculation data.

[0010] The first corrected calculation data is corrected based on the non-smooth distribution portion to obtain the second corrected calculation data. The second corrected calculation data is then locally smoothed according to the shock wave position to obtain the corrected data of the target wing surface.

[0011] The upper and lower wing surface corrected data are merged to obtain the corrected airfoil pressure distribution data.

[0012] Optionally, the process of sorting the acquired airfoil pressure distribution test data of the aircraft to obtain sorted data includes:

[0013] Read the airfoil pressure distribution test data of the aircraft from a file or database;

[0014] The airfoil pressure distribution test data are sorted in ascending order based on the normalized chord coordinates of the airfoil to obtain the sorted data.

[0015] Optionally, determining the target location where the minimum value in each target chord coordinate lies, and rearranging the airfoil pressure distribution calculation data based on the target location to obtain rearranged data, includes:

[0016] Read the airfoil pressure distribution calculation data from the file or database, obtain the maximum and minimum values ​​of the normalized chord coordinates of the airfoil pressure distribution calculation data, and determine the position corresponding to the minimum value;

[0017] According to the target location, the preset array storing the airfoil pressure distribution calculation data is traversed, and the traversed airfoil pressure distribution calculation data is stored in a temporary array.

[0018] After the preset array has been traversed completely, the airfoil pressure distribution calculation data is rearranged based on the data stored in the temporary array to obtain rearranged data; wherein, the minimum value is located at the first position of the rearranged data.

[0019] Optionally, the step of decomposing the rearranged data to obtain pressure distribution calculation data for the upper wing surface and pressure distribution calculation data for the lower wing surface includes:

[0020] Determine the difference between the maximum value of the normalized chord coordinate of the calculated airfoil pressure distribution data and the target threshold data;

[0021] The pressure distribution calculation data of the airfoil is traversed in a forward direction based on the target position. The traversal ends when any airfoil pressure distribution calculation data is greater than the difference, so as to obtain the pressure distribution calculation data of the upper airfoil.

[0022] The airfoil pressure distribution calculation data is traversed in reverse according to the target position. The traversal ends when any airfoil pressure distribution calculation data is greater than the difference, so as to obtain the pressure distribution calculation data of the lower airfoil.

[0023] Optionally, obtaining the target pressure coefficient calculation value at the test point location of the target wing surface based on the pressure distribution calculation data of the target wing surface includes:

[0024] Iterate through the calculated pressure distribution data of the target wing surface and the test value data of the pressure distribution of the target wing surface corresponding to the test point location of the target wing surface;

[0025] Determine whether the test point on the target wing surface is located within the target range;

[0026] If so, the target pressure coefficient at the test point location of the target wing surface is calculated based on the target interval using a preset linear difference formula.

[0027] Optionally, expanding the pressure coefficient difference to obtain the expanded difference includes:

[0028] The maximum and minimum values ​​of the normalized chord coordinates calculated based on the airfoil pressure distribution are each extended by one data point on the coordinate endpoints of the pressure coefficient difference to obtain the extended difference.

[0029] Accordingly, the smoothing operation on the expanded difference to decompose the expanded difference into a smooth distribution part and a non-smooth distribution part includes:

[0030] The expanded difference is determined as the initial non-optical surplus.

[0031] Based on the initial non-smooth surplus, the pressure coefficient difference is subjected to Laplace smoothing to obtain a smooth distribution portion;

[0032] The initial non-smooth surplus is subtracted from the smooth pressure coefficient difference corresponding to the smooth distribution portion to obtain the non-smooth distribution portion.

[0033] Optionally, the step of correcting the airfoil pressure distribution calculation data based on the smooth distribution portion to obtain the first corrected calculation data includes:

[0034] The smooth pressure coefficient difference is interpolated using the radial basis function method to obtain the corrected pressure difference value;

[0035] The corrected pressure difference value is superimposed with the pressure distribution calculation data of the target wing surface to obtain the first corrected calculation data.

[0036] Optionally, the step of correcting the first corrected calculated data based on the non-smooth distribution portion to obtain the second corrected calculated data includes:

[0037] The non-smooth distribution portion of the first n-1 layers is obtained, and the non-smooth distribution portion of the first n-1 layers is interpolated to the calculation point position through a linear interpolation method to obtain the pressure coefficient difference of the calculation point position of the target wing surface.

[0038] The pressure coefficient difference corrected by the calculation point position of the target wing surface is superimposed with the first corrected calculation data to obtain the second corrected calculation data.

[0039] Optionally, the step of performing a local smoothing operation on the second corrected calculated data based on the shock wave position to obtain the corrected airfoil data of the target airfoil includes:

[0040] The shock wave location is determined by detecting the shock wave position in the pressure distribution test data of the target wing surface in the sorted data.

[0041] Lagrange interpolation correction is used to interpolate and correct several array points before and after the shock wave position in the second corrected calculation data to obtain the corrected data for the upper wing surface and the corrected data for the lower wing surface.

[0042] Optionally, merging the upper and lower wing surface corrected data in the corrected wing surface data to obtain the corrected airfoil pressure distribution data includes:

[0043] The corrected data for the upper and lower wing surfaces are sorted and merged in a counter-clockwise order to obtain the corrected airfoil pressure distribution data.

[0044] Secondly, this application discloses an airfoil pressure data correction device for aircraft, comprising:

[0045] The data rearrangement module is used to sort the acquired airfoil pressure distribution test data of the aircraft to obtain sorted data, determine the target position where the minimum value of each target chord coordinate is located, and rearrange the airfoil pressure distribution calculation data based on the target position to obtain rearranged data; the target chord coordinates are obtained by normalizing the airfoil pressure distribution calculation data of the aircraft.

[0046] The data decomposition module is used to decompose the rearranged data to obtain the pressure distribution calculation data of the upper wing surface and the pressure distribution calculation data of the lower wing surface.

[0047] The pressure coefficient difference determination module is used to obtain the target pressure coefficient calculation value at the test point location of the target wing surface based on the pressure distribution calculation data of the target wing surface, and to determine the pressure coefficient difference between the pressure distribution test data of the target wing surface in the sorted data and the target pressure coefficient calculation value; the target wing surface is the upper wing surface or the lower wing surface;

[0048] The correction module is used to expand the pressure coefficient difference to obtain an expanded difference, perform a smoothing operation on the expanded difference to decompose the expanded difference into a smooth distribution part and a non-smooth distribution part, and correct the airfoil pressure distribution calculation data based on the smooth distribution part to obtain the first corrected calculation data.

[0049] The wing surface correction data acquisition module is used to correct the first corrected calculation data based on the non-smooth distribution part to obtain the second corrected calculation data, and to perform a local smoothing operation on the second corrected calculation data according to the shock wave position to obtain the wing surface correction data of the target wing surface.

[0050] The data merging module is used to merge the upper and lower wing surface corrected data in the wing surface corrected data to obtain the corrected airfoil pressure distribution data.

[0051] Thirdly, this application discloses an electronic device, including:

[0052] Memory, used to store computer programs;

[0053] A processor is used to execute computer programs to implement the aforementioned method for correcting airfoil pressure data for aircraft.

[0054] Fourthly, this application discloses a computer-readable storage medium storing a computer program, which, when executed by a processor, implements the aforementioned method for correcting airfoil pressure data of an aircraft.

[0055] This application first sorts the acquired airfoil pressure distribution test data of the aircraft to obtain sorted data, determines the target position of the minimum value in each target chord coordinate, and rearranges the airfoil pressure distribution calculation data based on the target position to obtain rearranged data; the target chord coordinates are obtained by normalizing the airfoil pressure distribution calculation data of the aircraft; the rearranged data is decomposed to obtain pressure distribution calculation data for the upper wing surface and pressure distribution calculation data for the lower wing surface; the target pressure coefficient calculation value is obtained at the test point position of the target wing surface based on the pressure distribution calculation data of the target wing surface, and the pressure coefficient difference between the pressure distribution test data of the target wing surface in the sorted data and the target pressure coefficient calculation value is determined. The target airfoil is either the upper airfoil or the lower airfoil. The pressure coefficient difference is expanded to obtain an expanded difference. The expanded difference is then smoothed to decompose it into a smooth distribution part and a non-smooth distribution part. The airfoil pressure distribution calculation data is corrected based on the smooth distribution part to obtain first corrected calculation data. The first corrected calculation data is then corrected based on the non-smooth distribution part to obtain second corrected calculation data. The second corrected calculation data is then locally smoothed according to the shock wave position to obtain the airfoil surface correction data. The upper airfoil correction data and the lower airfoil correction data in the airfoil surface correction data are then merged to obtain the corrected airfoil pressure distribution data.

[0056] As can be seen, this application first uses linear interpolation to obtain pressure calculation data at the test point location. Then, it calculates the difference between the experimental and calculated pressure data at the test point location and expands the pressure difference data according to the difference pattern. Next, it decomposes the difference data into smooth and non-smooth distribution parts through multiple smoothing processes. The smooth part is corrected using a high-order interpolation method, and the non-smooth part is corrected using linear interpolation. Finally, the airfoil pressure data is locally smoothed by shock wave location detection to obtain airfoil pressure distribution correction data with high overall accuracy. In this way, this application can not only make up for the deficiencies of pressure test data and calculated data, but also maintain the high accuracy of wind tunnel test data and the high resolution characteristics of calculated data. It can optimize the pressure distribution of components such as wings and tail fins with high precision, improve the lift-to-drag ratio of the aircraft, and improve flight stability. It can also accurately correct complex flow regions such as shock waves and boundary layer separation, and solve the problem of differences between pressure test data and calculated data under high temperature and high pressure conditions. Attached Figure Description

[0057] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are only embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on the provided drawings without creative effort.

[0058] Figure 1 This is a flowchart of a method for correcting airfoil pressure data for an aircraft, as disclosed in this application.

[0059] Figure 2 This is a schematic diagram showing the distribution of pressure test data and calculation data for a two-dimensional airfoil disclosed in this application;

[0060] Figure 3 This is a schematic diagram of a process for calculating the maximum and minimum values ​​of data coordinates using a determined coefficient, as disclosed in this application.

[0061] Figure 4 This is a schematic diagram of rearranging computational data starting from the minimum value of the computational data coordinates disclosed in this application;

[0062] Figure 5 This is a schematic diagram of a method disclosed in this application for segmenting rearranged computational data to obtain upper and lower wing surface data;

[0063] Figure 6 This is a flowchart of the calculation process for an n-layer hierarchical interpolation method disclosed in this application;

[0064] Figure 7This is a schematic diagram of the correction results of calculation data for a two-dimensional airfoil disclosed in this application;

[0065] Figure 8 This is a schematic diagram of the structure of an airfoil pressure data correction device for an aircraft disclosed in this application;

[0066] Figure 9 This is a structural diagram of an electronic device disclosed in this application. Detailed Implementation

[0067] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0068] Currently, combining the advantages of both types of data in corrected data is an important approach to obtaining accurate aerodynamic loads. However, problems such as strong discontinuities in pressure distribution and insufficient experimental data pose significant challenges to data correction methods. Airfoil pressure distribution data corrected using a single interpolation method suffers from local jumps or insufficient accuracy, making it difficult to meet the data accuracy requirements for engineering applications. To address these technical problems, this invention discloses an airfoil pressure data correction method that can compensate for the deficiencies of pressure test data and calculated data, maintaining both the high accuracy of wind tunnel test data and the high resolution characteristics of calculated data. Furthermore, it can accurately correct complex flow regions such as shock waves and boundary layer separation, resolving the discrepancies between pressure test data and calculated data under high temperature and high pressure conditions.

[0069] See Figure 1 As shown in the figure, an embodiment of the present invention discloses a method for correcting airfoil pressure data of an aircraft, including:

[0070] Step S11: Sort the obtained airfoil pressure distribution test data of the aircraft to obtain sorted data, determine the target position where the minimum value of each target chord coordinate is located, and rearrange the airfoil pressure distribution calculation data based on the target position to obtain rearranged data; the target chord coordinates are obtained by normalizing the airfoil pressure distribution calculation data of the aircraft.

[0071] In this embodiment, the airfoil pressure distribution test data of the aircraft is first read from a file or database; the airfoil pressure distribution test data is then sorted in ascending order based on the normalized chord coordinates of the airfoil to obtain the sorted data. In a specific embodiment, based on the two-dimensional airfoil flow state, the airfoil pressure distribution test data is read from a file or database, and 53 test data points of the upper surface pressure of the two-dimensional airfoil are stored in the upper surface pressure distribution data array. , In, among them, This indicates the number of test points on the upper wing surface. The 48 test data points for the lower wing surface pressure are stored in the lower wing surface pressure distribution data array. , middle. This indicates the number of test points on the lower airfoil surface. Where x is the normalized chordal coordinate of the airfoil. Let denoted as σexp, where σupper represents the pressure coefficient at the airfoil surface location, σn represents the number of pressure distribution data points, and the subscripts exp, upper, and low represent the lower airfoil surface. Subsequently, to facilitate rapid positioning and interpolation in later steps, the airfoil pressure distribution test data obtained in step one are sorted in ascending order along the x-coordinate. Sorting can be achieved using any sorting algorithm; however, given the small amount of data involved, efficiency is not a concern, and the classic bubble sort algorithm is used. In other words, this application employs the classic bubble sort algorithm to sort the acquired two-dimensional airfoil pressure distribution test data in ascending order along the x-coordinate.

[0072] Next, airfoil pressure distribution calculation data is read from a file or database. The maximum and minimum values ​​of the normalized chord coordinates of the airfoil pressure distribution calculation data are obtained, and the position corresponding to the minimum value is determined. Based on the target position, a preset array storing the airfoil pressure distribution calculation data is traversed, and the traversed airfoil pressure distribution calculation data is stored in a temporary array. After the preset array has been completely traversed, the airfoil pressure distribution calculation data is rearranged based on the data stored in the temporary array to obtain rearranged data; wherein the minimum value is located at the beginning of the rearranged data. In this process, based on the two-dimensional airfoil flow state, two-dimensional airfoil pressure distribution calculation data is read from a specified file, and 304 pressure calculation data are stored in an array. , Among them This indicates the amount of data to be calculated. The subscript `calc` indicates a calculation, and `x` is the normalized chord coordinate of the airfoil. Let be the pressure coefficient at the airfoil surface location, and n be the number of pressure distribution data points. The distribution of airfoil pressure test data and calculated data is as follows: Figure 2(Where, the horizontal axis is the normalized coordinate of the chordal position of the airfoil structure (where x is the chordal position and c is the chordal length, and here the chordal length of the airfoil is c=1), and the vertical axis is the pressure coefficient of the upper and lower airfoil positions corresponding to the normalized horizontal axis of the airfoil structure.)

[0073] When obtaining the maximum and minimum values ​​of the normalized coordinate x in the two-dimensional airfoil pressure distribution calculation data, this application iterates through the array for the obtained two-dimensional pressure distribution calculation data. Compare all elements in the array step by step and update the current maximum value. Finally, the maximum value in the array is found. Similarly, the array is traversed. Compare and update the current minimum value step by step among all elements in the set. Finally, the minimum value in the array is found, and its position is recorded as ii. The specific process for finding and calculating the maximum and minimum values ​​and their indices is as follows: Figure 3 As shown. Then rearrange the two-dimensional airfoil pressure distribution calculation data. Iterate through the array starting from position ii, where the minimum value is located. , And store each array traversed into a temporary array. , If traversing to the array... , The last value is then iterated from the beginning of the array until the position ii-1 is reached. This can be expressed as:

[0074] ;

[0075] .

[0076] After the traversal is complete, the pressure distribution calculation data is updated using the data from the temporary array. A schematic diagram of the rearranged two-dimensional airfoil pressure distribution data is shown below. Figure 4 As shown. Implementation Its position in the array is 1. The update expression can be represented as:

[0077] .

[0078] Step S12: Decompose the rearranged data to obtain the pressure distribution calculation data of the upper wing surface and the pressure distribution calculation data of the lower wing surface.

[0079] In this embodiment, the airfoil pressure distribution calculation data is decomposed into two parts: upper and lower airfoil surfaces. Based on the obtained rearranged pressure distribution calculation data, the pressure calculation data for the upper and lower airfoil surfaces are obtained through forward and reverse traversal, respectively, as follows: Figure 5As shown. Airfoil pressure distribution calculation data is typically sorted for better visualization, either clockwise or counter-clockwise. Here, it is assumed that the original airfoil pressure distribution data was sorted clockwise. The difference between the maximum value of the normalized chord coordinates of the airfoil pressure distribution calculation data and the target threshold data is determined. The airfoil pressure distribution calculation data is traversed forward according to the target position, ending the traversal when any airfoil pressure distribution calculation data is greater than the difference, to obtain the pressure distribution calculation data for the upper airfoil surface. The airfoil pressure distribution calculation data is then traversed backward according to the target position, ending the traversal when any airfoil pressure distribution calculation data is greater than the difference, to obtain the pressure distribution calculation data for the lower airfoil surface. Specifically, from... Starting from the position, traverse the loop array in the forward direction. ,when At that time, exit the traversal and calculate the number of data points for the upper wing pressure coefficient. The calculated pressure distribution data on the upper wing surface can be expressed as:

[0080] ;

[0081] ;

[0082] from Starting from the position, traverse the loop array in reverse order. ,when At that time, exit the traversal and calculate the number of data points for the lower wing pressure coefficient. The calculated pressure distribution data on the lower wing surface can be expressed as:

[0083] ;

[0084] ;

[0085] in, This is used to filter the dense pressure coefficient calculation data at the trailing edge of the airfoil to avoid problems in subsequent interpolation and positioning searches. A relatively small value is usually chosen; here, 0.01 is selected. The -1 in the sequence number indicates a reverse loop.

[0086] Step S13: Obtain the target pressure coefficient calculation value at the test point location of the target wing surface based on the pressure distribution calculation data of the target wing surface, and determine the pressure coefficient difference between the pressure distribution test data of the target wing surface and the target pressure coefficient calculation value in the sorted data; the target wing surface is the upper wing surface or the lower wing surface.

[0087] In this embodiment, after obtaining the pressure distribution calculation data for the upper and lower wing surfaces, the system iterates through the pressure distribution calculation data for either the upper or lower wing surface, along with the test value data of the target wing surface corresponding to the test point location. It then determines whether the test point of the target wing surface is located within the target interval. If so, the target pressure coefficient calculation value for the test point location of the target wing surface is calculated based on the target interval using a preset linear difference formula. The upper wing surface calculation data is then used to input... Linear interpolation of the array yields the calculated pressure coefficient at the corresponding test point on the upper wing surface. The subscript `calc_upper_linear` indicates the upper wing surface calculation data at the linear interpolation alignment position. For example, the calculated pressure coefficient values ​​at the upper wing surface test point positions are obtained by linearly interpolating 151 calculated pressure coefficient data points to 53 test point positions. First, iterate through the coordinates of the experimental pressure coefficient values ​​on the upper wing surface. and calculated coordinates The value of the parameter r is used to determine the outcome. and interval Positional relationship:

[0088] ;

[0089] When r≤0, it indicates In the interval Within, then calculate the interval according to the following linear interpolation formula. Internal calculated pressure coefficient Interpolation:

[0090] ;

[0091] in, It is the coordinate of the test point after linear interpolation of the calculated point. The pressure coefficient value at the location, j takes the value of , representing the j-th element of the test coordinates after pressure coefficient interpolation. It calculates the coordinates of the point. The pressure difference value at the location, i takes the value of The subscripts i and i+1 represent the i-th and i+1-th elements of the pressure coefficient calculation data coordinates, respectively. The number of pressure distribution calculation data points after linear interpolation alignment. .

[0092] Next, the difference between the experimental and calculated values ​​of the pressure distribution on the target wing surface is calculated. Taking the upper wing surface as an example, based on the pressure coefficient calculated at the test point obtained by linear interpolation, the difference between the experimental and calculated values ​​after aligning the upper wing surface is expressed as follows:

[0093] ;

[0094] in, This represents the difference between the pressure test value and the calculated value, and the number of pressure difference distributions after linear interpolation alignment. .

[0095] Step S14: Expand the pressure coefficient difference to obtain an expanded difference. Perform a smoothing operation on the expanded difference to decompose it into a smooth distribution part and a non-smooth distribution part. Correct the airfoil pressure distribution calculation data based on the smooth distribution part to obtain the first corrected calculation data.

[0096] In this embodiment, after obtaining the pressure coefficient difference, the test point locations for the pressure coefficient difference are expanded. Based on the maximum and minimum values ​​of the normalized chord coordinates of the airfoil pressure distribution calculation data, one data point is added to each of the coordinate endpoints of the pressure coefficient difference to obtain the expanded difference. To ensure consistency between the experimental and calculated spatial distributions, the test point data is expanded by extrapolation, combining the patterns of the pressure distribution calculation data and the experimental data, to maintain consistency between the coordinate ranges of the pressure coefficient difference and the calculated data. The minimum value of the obtained calculated data coordinates is then used as the basis for the expansion. and maximum value For the pressure coefficient difference array , Each coordinate endpoint is extended by one data point 0 and , The expanded coordinate endpoint pressure coefficient difference is obtained as follows:

[0097] ;

[0098] ;

[0099] ;

[0100] ;

[0101] in The distribution of pressure coefficient differences at the expanded intermediate locations is as follows:

[0102] ;

[0103] The subscript expand_upper indicates the upper wing surface data after the pressure difference has been expanded.

[0104] Next, a smoothing operation is performed on the expanded difference to decompose it into a smooth distribution portion and a non-smooth distribution portion. In this process, the expanded difference is determined as the initial non-smooth difference; based on the initial non-smooth difference, the pressure coefficient difference is smoothed using Laplace smoothing to obtain the smooth distribution portion; the smooth pressure coefficient difference corresponding to the smooth distribution portion is subtracted from the initial non-smooth difference to obtain the non-smooth distribution portion. To facilitate multi-layer interpolation iterative updates of interpolation data, taking the upper wing surface as an example, firstly, the initial pressure difference after expansion on the upper wing surface is assigned the initial non-smooth difference, and the upper wing surface pressure coefficient is assigned to the initial pressure coefficient:

[0105] ;

[0106] ;

[0107] In this context, the superscript 0 indicates the 0th iteration, the subscript calc_upper_interp represents the calculated data for the upper wing interpolation correction, and the subscript remain_upper represents the non-smooth distribution of the pressure difference on the upper wing. Secondly, before smoothing each interpolation iteration, the non-smooth portion of the previous layer is assigned to the smoothing difference portion, and the initial non-smooth difference after the upper wing expansion is assigned to the initial smoothing difference portion.

[0108] ;

[0109] The subscript `smooth_upper` represents the smoothed distribution of the pressure difference on the upper wing surface. Before each interpolation iteration, the expanded pressure coefficient difference on the upper wing surface is smoothed multiple times using a smoothing method. For each iteration of smoothing, the coordinate endpoints are preserved. , and , The value remains unchanged:

[0110] ;

[0111] ;

[0112] ;

[0113] ;

[0114] The superscripts 0 and 1 represent the 0th and 1st layer iterations, respectively, for the pressure coefficient difference array at the intermediate position. , Smoothing was achieved by using the k-th order Laplace smoothing method to obtain a smooth distribution. , The first layer smoothing formula is as follows:

[0115] ;

[0116] ;

[0117] in, It is an array , The pressure coefficient difference value of the i-th element. The coordinates of the test point location after smoothing the pressure coefficient difference of the upper wing surface extension. The value at the position, with the superscript 1 indicating the first level of interpolation iterative smoothing processing, and the subscript... This represents the i-th element of the coordinates of the test pressure difference after smoothing treatment. k is the iteration number of the smoothing treatment; to achieve a better smoothing effect, k=10. λ is the smoothing relaxation factor, which is set to 1 / 2. The non-smooth distribution of the first layer is obtained by subtracting the pressure coefficient difference of the first layer from the initial non-smooth pressure coefficient difference, as expressed below:

[0118] ;

[0119] In one specific embodiment, the difference in quantity between the smooth pressure distribution portion and the non-smooth pressure distribution portion of the upper wing surface is . =55.

[0120] In this embodiment, the smooth pressure coefficient difference is interpolated using the radial basis function method to obtain a corrected pressure difference value. This corrected pressure difference value is then superimposed with the pressure distribution calculation data of the target wing surface to obtain the first corrected calculation data. In other words, the smooth distribution portion of the upper wing surface pressure coefficient difference is interpolated at a higher order to the pressure coefficient calculation point, thus correcting the pressure coefficient calculation data. This corrects the pressure coefficient calculation data based on the obtained smooth distribution portion of the pressure coefficient difference. , First, the classic radial basis function (RBF) method is used to interpolate the smooth pressure coefficient difference to obtain the corrected pressure difference value for the first interpolation layer:

[0121] ;

[0122] in, This represents the pressure difference at coordinate x after RBF interpolation, where x is the location where the pressure coefficient is calculated. The coordinates, with the subscript `smooth_upper_RBF` indicating that the smooth portion of the upper wing pressure coefficient difference is interpolated using RBF interpolation. N represents the number of radial basis function base points, taking values ​​of... This refers to the location of the radial basis function base point, i.e., the coordinates of the expanded pressure differential data. The distance between two points is usually calculated using Euclidean distance. These are the interpolation coefficients. For the interpolation basis functions, this application uses Wendland's compact basis function C2 (that is, the Wendland function that is globally twice differentiable) to obtain them:

[0123] ;

[0124] in, R is the support radius of the basis function, with a value of 0.5, which is used to control the influence area of ​​each basis point.

[0125] Then, the difference after RBF interpolation is superimposed with the original pressure coefficient calculation data to obtain the first corrected pressure coefficient calculation data, as shown in the following expression:

[0126] ;

[0127] Wherein, the superscript 0 indicates the calculated value of the initial pressure coefficient of the 0th layer, and the superscript 1 indicates the light surplus of the 1st layer interpolated to the calculation point position via RBF.

[0128] After smoothing the pressure coefficient difference on the upper wing surface multiple times and then interpolating it to the pressure coefficient calculation point, the corrected pressure coefficient calculation data for the first n-1 layers is obtained. The process involves smoothing the expanded pressure coefficient difference on the upper wing surface, decomposing the pressure coefficient difference into smooth and non-smooth distribution parts, and then iterating the smooth distribution part of the pressure coefficient difference to the pressure coefficient calculation point to correct the pressure coefficient calculation data. This process is repeated n-1 times iteratively. In one specific embodiment, 10 smoothing processes are performed before interpolating each layer of RBF. For the (n-1)th layer, the non-smooth distribution part of the n-1th layer is obtained.

[0129] ;

[0130] The number of smooth and non-smooth parts in each layer of the upper wing surface is... =55, RBF interpolation is performed on the non-smooth distribution of the (n-1)th layer to obtain the pressure coefficient correction difference of the (n-1)th layer. The pressure difference corrected by RBF interpolation was then compared with the pressure coefficients calculated for the first n-2 layers. The pressure coefficient correction result for the (n-1)th layer is obtained by superposition:

[0131] ;

[0132] To achieve a better pressure coefficient correction effect, n=11 is chosen, where the superscripts n-1 and n-2 represent the (n-1)th and (n-2)th iterations, respectively. The number of higher-order interpolation corrections for the smooth pressure distribution on the upper wing surface is... =151.

[0133] Step S15: Correct the first corrected calculation data based on the non-smooth distribution portion to obtain the second corrected calculation data. Perform a local smoothing operation on the second corrected calculation data according to the shock wave position to obtain the corrected data of the target wing surface.

[0134] In this embodiment, the first n-1 layers of non-smooth distribution are obtained. A linear interpolation method is used to interpolate the first n-1 layers of non-smooth distribution to the calculation point position to obtain the pressure coefficient difference corrected for the calculation point position of the target wing surface. This corrected pressure coefficient difference is then superimposed with the first corrected calculation data to obtain the second corrected calculation data. The non-smooth distribution portion of the upper wing surface pressure coefficient difference is linearly interpolated to the pressure coefficient calculation point position to further correct the pressure coefficient calculation data. This is based on the array of the n-1 layers of non-smooth distribution in the smooth decomposition. , Using the linear interpolation method in step seven, the pressure coefficient difference corrected for the calculation point position on the upper wing surface is obtained by interpolating to the calculation point position:

[0135] ;

[0136] The subscript `remain_upper_linear` indicates the location of the non-smooth distribution portion of the linear interpolation to the calculation point. The coordinates of the non-smooth pressure distribution are linearly interpolated to the location of the calculation point. The pressure difference value at the location, i takes the value of This represents the i-th element of the test coordinates after pressure coefficient interpolation. It calculates the coordinates of the point. The pressure difference value at the point, j takes the value of The subscripts j and j+1 represent the j-th and j+1-th elements of the coordinate data for calculating the pressure difference, respectively.

[0137] Then, the pressure coefficient difference corrected by the non-smooth distribution interpolation is superimposed with the correction results of the first n-1 layers to obtain the pressure coefficient calculation data after n layers correction, as shown in the following expression:

[0138] ;

[0139] In this context, the superscripts n and n-1 represent the interpolated pressure coefficient calculation data for the n=11th and n-1=10th layers, respectively.

[0140] Next, shock wave location detection is performed on the pressure distribution test data of the target wing surface in the sorted data to determine the shock wave location. Then, Lagrange interpolation correction is applied to several array points before and after the shock wave location in the second corrected calculated data to obtain corrected data for the upper and lower wing surfaces. Taking the upper wing surface as an example, the upper wing surface pressure distribution data array is traversed. , The intermediate distribution after removing the first and last coordinate endpoints is calculated using the forward difference method. Pressure gradient values ​​at each point:

[0141] ;

[0142] in Indicates position The pressure gradient at point i, where i takes values ​​from 2, ... ,when At that time, shock wave data points are detected based on the absolute value of the pressure gradient ratio between any two adjacent points. Time stamp number Points , The location of the shock wave, where To set the shock wave detection threshold, in one specific embodiment, the shock wave location data coordinates are:

[0143] .

[0144] Next, the pressure coefficient correction data of the upper wing surface at the shock wave location is locally smoothed. This applies to the areas before and after the shock wave location. array points , The computational data is reconstructed using classic Lagrange interpolation:

[0145] ;

[0146] Among them, take =3, the superscript lg indicates the Lagrange interpolation correction. Here are the expressions for the Lagrange basis functions:

[0147] ;

[0148] After locally smoothing the pressure coefficient calculation data after correction for n=11 layers (the second corrected calculation data), the final corrected upper wing pressure coefficient array is obtained. , The subscript `calc_upper_Modified` indicates the calculated data for the final upper wing interpolation correction. The complete calculation flow for the n-layer hierarchical interpolation method is as follows:Figure 6 As shown.

[0149] Step S16: Merge the upper and lower wing surface corrected data in the corrected wing surface data to obtain the corrected airfoil pressure distribution data.

[0150] In this embodiment, although the above process is described using the upper wing surface as an example, the pressure coefficient data correction process for the lower wing surface is the same as that for the upper wing surface. After obtaining the corrected data for both the upper and lower wing surfaces from the corrected data, the corrected data for the upper and lower wing surfaces is sorted and merged in a counter-clockwise order to obtain the corrected airfoil pressure distribution data. Specifically, the pressure coefficient array is sorted and synthesized using a counter-clockwise order to achieve interpolation correction. , The subscript calc_Modified indicates the calculated data for the final lower wing surface interpolation correction.

[0151] Starting from the position of the maximum value of the corrected lower wing surface horizontal coordinate, the pressure coefficient of the lower wing surface is stored in reverse order, which can be expressed as:

[0152] ;

[0153] ;

[0154] Then, starting from the position of the minimum value of the corrected upper wing surface horizontal coordinate, we continue to store the corrected pressure coefficient of the upper wing surface by traversing in a forward direction, which can be expressed as:

[0155] ;

[0156] ;

[0157] Finally, the arrays of original experimental upper and lower wing surface data, calculated data, and interpolated corrected calculated data are saved together to a results file. The visualization results are shown below. Figure 7As shown (where the horizontal axis represents the normalized coordinates of the chordal position of the airfoil structure (x is the chordal position, c is the chordal length, here the airfoil chordal length c=1), and the vertical axis represents the pressure coefficients of the upper and lower airfoil positions corresponding to the normalized horizontal axis of the airfoil structure). It should be noted that the above example is only one case of the layered interpolation method. The pressure data correction method based on layered interpolation provided by this invention can be extended to the pressure data correction of various airfoils according to the airfoil flow state. All extensions or improvements made based on the principles and methods of this invention are within the protection scope of this invention. This allows for highly accurate optimization of the pressure distribution of components such as wings and tail fins, improving the lift-to-drag ratio of the aircraft and enhancing flight stability. Furthermore, addressing the interpolation difficulties faced by traditional interpolation methods, such as the sparsity of wind tunnel test data and the strong discontinuities in pressure distribution, this invention can accurately correct complex flow regions such as shock waves and boundary layer separation, solving the problem of discrepancies between pressure test data and calculated data under high temperature and high pressure conditions.

[0158] As can be seen, this application first uses linear interpolation to obtain pressure calculation data at the test point location. Then, it calculates the difference between the experimental and calculated pressure data at the test point location and expands the pressure difference data according to the difference pattern. Next, it decomposes the difference data into smooth and non-smooth distribution parts through multiple smoothing processes. The smooth part is corrected using a high-order interpolation method, and the non-smooth part is corrected using linear interpolation. Finally, the airfoil pressure data is locally smoothed by shock wave location detection to obtain airfoil pressure distribution correction data with high overall accuracy. In this way, this application can not only make up for the deficiencies of pressure test data and calculated data, but also maintain the high accuracy of wind tunnel test data and the high resolution characteristics of calculated data. It can optimize the pressure distribution of components such as wings and tail fins with high precision, improve the lift-to-drag ratio of the aircraft, and improve flight stability. It can also accurately correct complex flow regions such as shock waves and boundary layer separation, and solve the problem of differences between pressure test data and calculated data under high temperature and high pressure conditions.

[0159] See Figure 8 As shown, an embodiment of the present invention discloses an airfoil pressure data correction device, comprising:

[0160] The data rearrangement module 11 is used to sort the acquired airfoil pressure distribution test data of the aircraft to obtain sorted data, determine the target position where the minimum value of each target chord coordinate is located, and rearrange the airfoil pressure distribution calculation data based on the target position to obtain rearranged data; the target chord coordinates are obtained by normalizing the airfoil pressure distribution calculation data of the aircraft.

[0161] The data decomposition module 12 is used to decompose the rearranged data to obtain the pressure distribution calculation data of the upper wing surface and the pressure distribution calculation data of the lower wing surface.

[0162] The pressure coefficient difference determination module 13 is used to obtain the target pressure coefficient calculation value of the test point location of the target wing surface based on the pressure distribution calculation data of the target wing surface, and to determine the pressure coefficient difference between the pressure distribution test data of the target wing surface in the sorted data and the target pressure coefficient calculation value; the target wing surface is the upper wing surface or the lower wing surface;

[0163] The correction module 14 is used to expand the pressure coefficient difference to obtain an expanded difference, perform a smoothing operation on the expanded difference to decompose the expanded difference into a smooth distribution part and a non-smooth distribution part, and correct the airfoil pressure distribution calculation data based on the smooth distribution part to obtain the first corrected calculation data.

[0164] The wing surface correction data acquisition module 15 is used to correct the first corrected calculation data based on the non-smooth distribution part to obtain the second corrected calculation data, and to perform a local smoothing operation on the second corrected calculation data according to the shock wave position to obtain the wing surface correction data of the target wing surface.

[0165] The data merging module 16 is used to merge the upper wing surface corrected data and the lower wing surface corrected data in the wing surface corrected data to obtain the corrected airfoil pressure distribution data.

[0166] Since the embodiments of the device part correspond to the embodiments described above, please refer to the embodiments described in the method part for the embodiments of the device part, and will not be repeated here.

[0167] As can be seen, this application first uses linear interpolation to obtain pressure calculation data at the test point location. Then, it calculates the difference between the experimental and calculated pressure data at the test point location and expands the pressure difference data according to the difference pattern. Next, it decomposes the difference data into smooth and non-smooth distribution parts through multiple smoothing processes. The smooth part is corrected using a high-order interpolation method, and the non-smooth part is corrected using linear interpolation. Finally, the airfoil pressure data is locally smoothed by shock wave location detection to obtain airfoil pressure distribution correction data with high overall accuracy. In this way, this application can not only make up for the deficiencies of pressure test data and calculated data, but also maintain the high accuracy of wind tunnel test data and the high resolution characteristics of calculated data. It can optimize the pressure distribution of components such as wings and tail fins with high precision, improve the lift-to-drag ratio of the aircraft, and improve flight stability. It can also accurately correct complex flow regions such as shock waves and boundary layer separation, and solve the problem of differences between pressure test data and calculated data under high temperature and high pressure conditions.

[0168] Furthermore, embodiments of this application also disclose an electronic device, Figure 9 This is a structural diagram of an electronic device 20 according to an exemplary embodiment. The content of the diagram should not be construed as limiting the scope of this application.

[0169] Figure 9 This is a schematic diagram of the structure of an electronic device 20 provided in an embodiment of this application. Specifically, the electronic device 20 may include: at least one processor 21, at least one memory 22, a power supply 23, a communication interface 24, an input / output interface 25, and a communication bus 26. The memory 22 stores a computer program, which is loaded and executed by the processor 21 to implement the relevant steps in the aircraft airfoil pressure data correction method disclosed in any of the foregoing embodiments. Alternatively, the electronic device 20 in this embodiment may specifically be an electronic computer.

[0170] In this embodiment, the power supply 23 is used to provide operating voltage for each hardware device on the electronic device 20; the communication interface 24 can create a data transmission channel between the electronic device 20 and external devices, and the communication protocol it follows can be any communication protocol applicable to the technical solution of this application, and is not specifically limited here; the input / output interface 25 is used to acquire external input data or output data to the outside world, and its specific interface type can be selected according to specific application needs, and is not specifically limited here.

[0171] In addition, the memory 22, as a carrier for resource storage, can be a read-only memory, random access memory, disk or optical disk, etc. The resources stored thereon can include operating system 221, computer program 222, etc., and the storage method can be temporary storage or permanent storage.

[0172] The operating system 221 is used to manage and control the various hardware devices on the electronic device 20 and the computer program 222, which may be Windows Server, Netware, Unix, Linux, etc. In addition to including a computer program capable of performing the aircraft airfoil pressure data correction method executed by the electronic device 20 as disclosed in any of the foregoing embodiments, the computer program 222 may further include computer programs capable of performing other specific tasks.

[0173] Furthermore, this application also discloses a computer-readable storage medium for storing a computer program; wherein, when the computer program is executed by a processor, it implements the aforementioned method for correcting airfoil pressure data of an aircraft. Specific steps of this method can be found in the corresponding content disclosed in the foregoing embodiments, and will not be repeated here.

[0174] The various embodiments in this specification are described in a progressive manner, with each embodiment focusing on its differences from other embodiments. Similar or identical parts between embodiments can be referred to interchangeably. For the apparatus disclosed in the embodiments, since it corresponds to the method disclosed in the embodiments, the description is relatively simple; relevant parts can be referred to in the method section.

[0175] Those skilled in the art will further recognize that the units and algorithm steps of the various examples described in conjunction with the embodiments disclosed herein can be implemented in electronic hardware, computer software, or a combination of both. To clearly illustrate the interchangeability of hardware and software, the components and steps of the various examples have been generally described in terms of functionality in the foregoing description. Whether these functions are implemented in hardware or software depends on the specific application and design constraints of the technical solution. Those skilled in the art can use different methods to implement the described functions for each specific application, but such implementation should not be considered beyond the scope of this application.

[0176] The steps of the methods or algorithms described in conjunction with the embodiments disclosed herein can be implemented directly by hardware, a software module executed by a processor, or a combination of both. The software module can be located in random access memory (RAM), main memory, read-only memory (ROM), electrically programmable ROM, electrically erasable programmable ROM, registers, hard disk, removable disk, CD-ROM, or any other form of storage medium known in the art.

[0177] Finally, it should be noted that in this document, relational terms such as "first" and "second" are used only to distinguish one entity or operation from another, and do not necessarily require or imply any such actual relationship or order between these entities or operations. Furthermore, the terms "comprising," "including," or any other variations thereof are intended to cover non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements includes not only those elements but also other elements not expressly listed, or elements inherent to such a process, method, article, or apparatus. Without further limitations, an element defined by the phrase "comprising one..." does not exclude the presence of other identical elements in the process, method, article, or apparatus that includes said element.

[0178] The technical solutions provided in this application have been described in detail above. Specific examples have been used to illustrate the principles and implementation methods of this application. The descriptions of the above embodiments are only for the purpose of helping to understand the methods and core ideas of this application. At the same time, for those skilled in the art, there will be changes in the specific implementation methods and application scope based on the ideas of this application. Therefore, the content of this specification should not be construed as a limitation of this application.

Claims

1. A method for correcting airfoil pressure data of an aircraft, characterized in that, include: The acquired airfoil pressure distribution test data of the aircraft is sorted to obtain sorted data. The target position of the minimum value in each target chord coordinate is determined. Based on the target position, the airfoil pressure distribution calculation data is rearranged to obtain rearranged data. The target chord coordinates are obtained by normalizing the airfoil pressure distribution calculation data of the aircraft. The rearranged data is decomposed to obtain the pressure distribution calculation data of the upper wing surface and the pressure distribution calculation data of the lower wing surface; The calculated value of the target pressure coefficient at the test point location of the target wing surface is obtained based on the pressure distribution calculation data of the target wing surface, and the pressure coefficient difference between the pressure distribution test data of the target wing surface and the calculated value of the target pressure coefficient in the sorted data is determined; the target wing surface is the upper wing surface or the lower wing surface; The pressure coefficient difference is extended to obtain an extended difference. The extended difference is then smoothed to decompose it into a smooth distribution part and a non-smooth distribution part. The airfoil pressure distribution calculation data is then corrected based on the smooth distribution part to obtain the first corrected calculation data. The first corrected calculation data is corrected based on the non-smooth distribution portion to obtain the second corrected calculation data. The second corrected calculation data is then locally smoothed according to the shock wave position to obtain the corrected data of the target wing surface. The upper and lower wing surface corrected data in the corrected wing surface data are merged to obtain the corrected airfoil pressure distribution data; The process of expanding the pressure coefficient difference to obtain the expanded difference includes: The maximum and minimum values ​​of the normalized chord coordinates calculated based on the airfoil pressure distribution are each extended by one data point on the coordinate endpoints of the pressure coefficient difference to obtain the extended difference. Accordingly, the smoothing operation on the expanded difference to decompose the expanded difference into a smooth distribution part and a non-smooth distribution part includes: The expanded difference is determined as the initial non-optical surplus. Based on the initial non-smooth surplus, the pressure coefficient difference is subjected to Laplace smoothing to obtain a smooth distribution portion; The initial non-smooth surplus is subtracted from the smooth pressure coefficient difference corresponding to the smooth distribution portion to obtain the non-smooth distribution portion.

2. The method for correcting airfoil pressure data of an aircraft according to claim 1, characterized in that, The obtained airfoil pressure distribution test data of the aircraft is sorted to obtain sorted data, including: Read the airfoil pressure distribution test data of the aircraft from a file or database; The airfoil pressure distribution test data are sorted in ascending order based on the normalized chord coordinates of the airfoil to obtain the sorted data.

3. The method for correcting airfoil pressure data of an aircraft according to claim 1, characterized in that, The process involves determining the target location where the minimum value is located in each target chord coordinate, and then rearranging the airfoil pressure distribution calculation data based on the target location to obtain rearranged data, including: Read the airfoil pressure distribution calculation data from the file or database, obtain the maximum and minimum values ​​of the normalized chord coordinates of the airfoil pressure distribution calculation data, and determine the position corresponding to the minimum value; According to the target location, the preset array storing the airfoil pressure distribution calculation data is traversed, and the traversed airfoil pressure distribution calculation data is stored in a temporary array. After the preset array has been traversed completely, the airfoil pressure distribution calculation data is rearranged based on the data stored in the temporary array to obtain rearranged data; wherein, the minimum value is located at the first position of the rearranged data.

4. The method for correcting airfoil pressure data of an aircraft according to claim 3, characterized in that, The step of decomposing the rearranged data to obtain pressure distribution calculation data for the upper wing surface and pressure distribution calculation data for the lower wing surface includes: Determine the difference between the maximum value of the normalized chord coordinate of the calculated airfoil pressure distribution data and the target threshold data; The pressure distribution calculation data of the airfoil is traversed in a forward direction based on the target position. The traversal ends when any airfoil pressure distribution calculation data is greater than the difference, so as to obtain the pressure distribution calculation data of the upper airfoil. The airfoil pressure distribution calculation data is traversed in reverse based on the target position. The traversal ends when any airfoil pressure distribution calculation data is greater than the difference, so as to obtain the pressure distribution calculation data of the lower airfoil.

5. The method for correcting airfoil pressure data of an aircraft according to claim 1, characterized in that, The step of obtaining the target pressure coefficient calculation value at the test point location of the target wing surface based on the pressure distribution calculation data of the target wing surface includes: Iterate through the calculated pressure distribution data of the target wing surface and the test value data of the pressure distribution of the target wing surface corresponding to the test point location of the target wing surface; Determine whether the test point on the target wing surface is located within the target range; If so, the target pressure coefficient at the test point location of the target wing surface is calculated based on the target interval using a preset linear difference formula.

6. The method for correcting airfoil pressure data of an aircraft according to claim 1, characterized in that, The step of correcting the airfoil pressure distribution calculation data based on the smooth distribution to obtain the first corrected calculation data includes: The smooth pressure coefficient difference is interpolated using the radial basis function method to obtain the corrected pressure difference value; The corrected pressure difference value is superimposed with the pressure distribution calculation data of the target wing surface to obtain the first corrected calculation data.

7. The method for correcting airfoil pressure data of an aircraft according to claim 1, characterized in that, The step of correcting the first corrected calculated data based on the non-smooth distribution portion to obtain the second corrected calculated data includes: The non-smooth distribution portion of the first n-1 layers is obtained, and the non-smooth distribution portion of the first n-1 layers is interpolated to the calculation point position through a linear interpolation method to obtain the pressure coefficient difference of the calculation point position of the target wing surface. The pressure coefficient difference corrected by the calculation point position of the target wing surface is superimposed with the first corrected calculation data to obtain the second corrected calculation data.

8. The method for correcting airfoil pressure data of an aircraft according to any one of claims 1 to 7, characterized in that, The step of performing a local smoothing operation on the second corrected calculated data based on the shock wave position to obtain the corrected airfoil data for the target airfoil includes: The shock wave location is determined by detecting the shock wave position in the pressure distribution test data of the target wing surface in the sorted data. Lagrange interpolation correction is used to interpolate and correct several array points before and after the shock wave position in the second corrected calculation data to obtain the corrected data for the upper wing surface and the corrected data for the lower wing surface.

9. The method for correcting airfoil pressure data of an aircraft according to claim 8, characterized in that, The process of merging the upper and lower wing surface corrected data in the corrected wing surface data to obtain the corrected airfoil pressure distribution data includes: The corrected data for the upper and lower wing surfaces are sorted and merged in a counter-clockwise order to obtain the corrected airfoil pressure distribution data.

10. A device for correcting airfoil pressure data of an aircraft, characterized in that, include: The data rearrangement module is used to sort the acquired airfoil pressure distribution test data of the aircraft to obtain sorted data, determine the target position where the minimum value of each target chord coordinate is located, and rearrange the airfoil pressure distribution calculation data based on the target position to obtain rearranged data; the target chord coordinates are obtained by normalizing the airfoil pressure distribution calculation data of the aircraft. The data decomposition module is used to decompose the rearranged data to obtain the pressure distribution calculation data of the upper wing surface and the pressure distribution calculation data of the lower wing surface. The pressure coefficient difference determination module is used to obtain the target pressure coefficient calculation value at the test point location of the target wing surface based on the pressure distribution calculation data of the target wing surface, and to determine the pressure coefficient difference between the pressure distribution test data of the target wing surface in the sorted data and the target pressure coefficient calculation value; the target wing surface is the upper wing surface or the lower wing surface; The correction module is used to expand the pressure coefficient difference to obtain an expanded difference, perform a smoothing operation on the expanded difference to decompose the expanded difference into a smooth distribution part and a non-smooth distribution part, and correct the airfoil pressure distribution calculation data based on the smooth distribution part to obtain the first corrected calculation data. The wing surface correction data acquisition module is used to correct the first corrected calculation data based on the non-smooth distribution part to obtain the second corrected calculation data, and to perform a local smoothing operation on the second corrected calculation data according to the shock wave position to obtain the wing surface correction data of the target wing surface. The data merging module is used to merge the upper wing surface corrected data and the lower wing surface corrected data in the wing surface corrected data to obtain the corrected airfoil pressure distribution data; The correction module is used to extend the coordinate endpoints of the pressure coefficient difference by one data point each, based on the maximum and minimum values ​​of the normalized chord coordinates of the airfoil pressure distribution calculation data, to obtain the extended difference. The expanded difference is determined as the initial non-optical surplus. Based on the initial non-smooth surplus, the pressure coefficient difference is smoothed using Laplace smoothing to obtain a smooth distribution portion; the initial non-smooth surplus is subtracted from the smooth pressure coefficient difference corresponding to the smooth distribution portion to obtain the non-smooth distribution portion.

11. An electronic device, characterized in that, include: Memory, used to store computer programs; A processor for executing a computer program to implement the steps of the aircraft airfoil pressure data correction method as described in any one of claims 1 to 9.

12. A computer-readable storage medium, characterized in that, A computer program is stored on a computer-readable storage medium, which, when executed by a processor, implements the steps of the airfoil pressure data correction method as described in any one of claims 1 to 9.

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