A method for processing aerodynamic data for a high camber wing

Abstract

The application belongs to the technical field of aircraft aerodynamic load design, and is an aerodynamic data processing method for a high-camber wing. Target spanwise position-based spanwise coordinate groups are extracted based on the aircraft body coordinate system, target spanwise data of the wing is input into numerical calculation software or data processing software, external environmental parameters are given, pressure coefficient distribution data of the target spanwise position is calculated and exported; through comparison and judgment with the airfoil leading edge point, coordinate point sets at different positions are obtained, and surface pressure coefficients in different coordinate point sets are extracted respectively, the surface pressure coefficients are sequentially interpolated into chordwise positions required for load calculation, and chordwise position data of the high-camber wing is obtained. The wing surface data is successfully processed through rearrangement of the coordinate points and division of the regions to distinguish the upper and lower surface points by different methods. The influence of the wing camber change is considered, and the design requirements in engineering can be met.

Description

Technical Field

[0001] This application belongs to the field of aircraft aerodynamic load design technology, and specifically relates to an aerodynamic data processing method for wings with large camber. Background Technology

[0002] As computing power continues to increase, the accuracy of numerical calculations of flow fields on aircraft surfaces continues to improve. However, the airfoil surface data extracted from the numerical calculation results cannot be directly used for aerodynamic load calculations. In particular, when the airfoil camber is large or the wing trailing edge control surface deflects, accurately re-interpolating the numerical calculation results is a necessary condition for their use in flight load design.

[0003] Currently, there are load data analysis techniques based on local mesh refinement methods. However, existing aerodynamic load analysis methods, when interpolating non-trapezoidal wings using re-mesh methods, suffer from low accuracy in the coordinate positions of calculated points, leading to significant errors in aerodynamic load calculations. Furthermore, as the wing trailing edge control surface deflection changes, the interpolation mesh must be re-refined. Therefore, there is a need to develop a more accurate data interpolation method that can adapt to more complex wing shapes and changes in control surface deflection. Summary of the Invention

[0004] The purpose of this application is to provide an aerodynamic data processing method for large-camber wings, so as to solve the problem that it is difficult to perform aerodynamic data processing on complex wing shapes in the prior art.

[0005] The technical solution of this application is: an aerodynamic data processing method for large-camber wings, comprising:

[0006] Determine the target spanwise position for load design, extract the spanwise coordinate set of the target spanwise position based on the aircraft body coordinate system, and mark each coordinate point in the spanwise coordinate set;

[0007] Input the wing target spanwise data into numerical calculation software or data processing software, give external environmental parameters, calculate and export the pressure coefficient distribution data of the target spanwise position;

[0008] The coordinate points in the pressure coefficient distribution data are compared with the airfoil leading edge points in sequence according to the marking order to determine the airfoil leading edge points in the spanwise coordinate group. Then, the points that belong to the upper surface of the airfoil are formed into the first set, and the coordinate points that belong to the lower surface of the airfoil are formed into the second set. The surface pressure coefficients of the coordinate points in the first set and the second set are extracted.

[0009] The surface pressure coefficients are interpolated sequentially to the chordal positions required for aerodynamic load calculation to obtain the chordal position data of the large-camber wing.

[0010] Preferably, the spanwise profile of the target spanwise position is parallel to or at a certain angle to the xoz plane of the aircraft body coordinate system.

[0011] Preferably, the specific calculation method for the pressure coefficient distribution data is as follows:

[0012] First, the pressure coefficient is calculated based on the external environmental parameters. Then, the three-dimensional coordinates and corresponding pressure coefficients at the spanwise profile are derived based on the target spanwise position. The basic form is (x,y,z,Cp).

[0013] Preferably, the specific method for determining the data within the first set and the second set is as follows:

[0014] The pressure coefficient distribution data in the basic form (x, y, z, Cp) are sorted in ascending order according to the x-coordinate;

[0015] The first point after sorting is taken as the leading edge coordinate point of the airfoil;

[0016] Points located in the first certain proportion of the sorted array are determined to be at the leading edge position of the airfoil at that spanwise position. If their z-coordinate is greater than the z-coordinate of the first point, then the point belongs to the upper surface of the airfoil at that spanwise position and is included in the first set; otherwise, the point belongs to the lower surface of the airfoil at that spanwise position and is included in the second set.

[0017] The remaining coordinate points are determined to be located in the mid-rear part of the airfoil in the spanwise direction. The first remaining coordinate point is selected according to the marked order, and its relative position to the last point on the upper surface that has already been determined is calculated:

[0018] ;

[0019] Among them, z up This is the last point on the upper surface that was previously determined to be complete, z down This is the last point on the lower surface that was previously determined to be complete;

[0020] K0 is the target threshold, if K i If the value is less than K0, then point i is considered to be on the upper surface of the airfoil and included in the first set; otherwise, point i is determined to be on the lower surface of the airfoil and included in the second set.

[0021] The position of the next coordinate point is determined according to the marked order until the remaining coordinate points have been determined.

[0022] Preferably, the specific method for obtaining the chordal position is as follows:

[0023] Calculate the relative chord length of the corresponding positions of the coordinate points in the first and second sets using the chord length calculation formula. The chord length calculation formula is as follows:

[0024] ;

[0025] Among them, xc i Let x be the relative chord length at point i. i Let x be the x-coordinate of point i, and x1 be the x-coordinate of the leading edge point. m The x-coordinate of the trailing edge point;

[0026] The relative distance is calculated using the relative chord length, and it is then determined whether the distance is less than a distance threshold. The determination formula is:

[0027] ;

[0028] If so, then the relative chord length is increased according to the distance threshold, as shown in the formula:

[0029] ;

[0030] If not, calculate the relative position of the corresponding coordinate point based on the load, perform linear interpolation on the arrays (xc, Cp) of the upper and lower surfaces respectively, and output the calculation results as required.

[0031] This application presents an aerodynamic data processing method for high-camber wings. It extracts a spanwise coordinate set based on the aircraft body coordinate system for the target spanwise position, inputs the data from the spanwise coordinate set into numerical calculation software or data processing software, and calculates and derives the pressure coefficient distribution data for the target spanwise position given external environmental parameters. By comparing this data with the airfoil leading edge point, it obtains sets of coordinate points at different locations and extracts the surface pressure coefficients from each set. These surface pressure coefficients are then interpolated to the chordal position required for load calculation, yielding the chordal pressure coefficients of the high-camber wing. By rearranging the coordinate points and using different methods to distinguish upper and lower surface points in different regions, the method successfully processes wing surface data (including high-camber wings). It considers the influence of wing camber variations and can meet design requirements in engineering. Attached Figure Description

[0032] To more clearly illustrate the technical solutions provided in this application, the accompanying drawings will be briefly described below. Obviously, the drawings described below are merely some embodiments of this application.

[0033] Figure 1 This is a schematic diagram of the aerodynamic data processing flow for the wing of this application;

[0034] Figure 2 This is a schematic diagram of a spanwise cross-section of the large-camber wing of this application;

[0035] Figure 3 This is a schematic diagram of the pressure coefficient re-interpolated from a spanwise section of the large-camber wing of this application. Detailed Implementation

[0036] 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.

[0037] An aerodynamic data processing method for high-camber wings, such as... Figure 1-2 As shown, it includes:

[0038] Step S100: Determine the target spanwise position for load design, extract the spanwise coordinate set of the target spanwise position based on the aircraft body coordinate system, and mark each coordinate point in the spanwise coordinate set. The spanwise profile of the target spanwise position is parallel to or at a certain angle to the xoz plane of the aircraft body coordinate system.

[0039] Step S200: Input the data in the spanwise coordinate group into numerical calculation software or data processing software, give the external environmental parameters, calculate and export the pressure coefficient distribution data of the target spanwise location.

[0040] Preferably, combined with Figure 3 The specific calculation method for the pressure coefficient distribution data is as follows:

[0041] Step S210: First, calculate the pressure coefficient Cp based on the external environmental parameters;

[0042] Step S220: Then, based on the target spanwise position, the three-dimensional coordinates and corresponding pressure coefficients at the spanwise profile are derived, with the basic form being (x, y, z, Cp).

[0043] Step S300: According to the marking order, the coordinate points in the pressure coefficient distribution data are compared with the airfoil leading edge points in turn to determine the airfoil leading edge points in the spanwise coordinate group. Then, the points that are determined to belong to the upper surface of the airfoil are formed into a first set, and the coordinate points that are determined to belong to the lower surface of the airfoil are formed into a second set. The surface pressure coefficients of the coordinate points in the first set and the second set are extracted.

[0044] Preferably, the specific method for determining the data within the first set and the second set is as follows:

[0045] Step S310: Sort the pressure coefficient distribution data in the basic form (x,y,z,Cp) according to the x-coordinate in ascending order;

[0046] Step S320: Take the first point after sorting as the coordinate point of the airfoil leading edge;

[0047] Step S330: The points located in the first certain proportion of the sorted array are determined to be at the leading edge position of the airfoil at the spanwise position, such as the first 20%, 30%, etc., and adjusted according to specific needs. If its z-coordinate is greater than the z-coordinate of the first point, then the point belongs to the upper surface of the airfoil at the spanwise position and is included in the first set; otherwise, the point belongs to the lower surface of the airfoil at the spanwise position and is included in the second set.

[0048] Step S340: Determine that the remaining coordinate points are located in the mid-rear part of the airfoil in the spanwise direction. Select the first remaining coordinate point according to the marking order and calculate its relative position with the last point on the upper surface that has already been determined:

[0049] ;

[0050] Among them, z up This is the last point on the upper surface that was previously determined to be complete, z down This is the last point on the lower surface that was previously determined to be complete;

[0051] Step S350, K0 is the target threshold, if K i If the value is less than K0, then point i is considered to be on the upper surface of the airfoil and included in the first set; otherwise, point i is determined to be on the lower surface of the airfoil and included in the second set.

[0052] Step S360: Determine the position of the next coordinate point according to the marked order, until the remaining coordinate points are determined.

[0053] Step S400: Interpolate the surface pressure coefficients sequentially to the chordal position required for load calculation to obtain the chordal position surface pressure data of the large-camber wing.

[0054] Preferably, the specific method for obtaining the chordal position is as follows:

[0055] Step S410: Calculate the relative chord length of the corresponding positions of the coordinate points in the first and second sets using the chord length calculation formula. The chord length calculation formula is:

[0056] ;

[0057] Among them, xc i Let x be the relative chord length at point i. i Let x be the x-coordinate of point i, and x1 be the x-coordinate of the leading edge point. m The x-coordinate of the trailing edge point;

[0058] Step S420: Calculate the relative distance using the relative chord length and determine if it is less than a distance threshold. The determination formula is:

[0059] ;

[0060] If so, then the relative chord length is increased according to the distance threshold, as shown in the formula:

[0061] ;

[0062] Step S430: If not, calculate the relative position of the corresponding coordinate point according to the load, perform linear interpolation on the array (xc, Cp) of the upper and lower surfaces respectively, and output the calculation results as required.

[0063] This application extracts the spanwise coordinate set of the target spanwise position based on the aircraft body coordinate system. The data from the spanwise coordinate set is input into numerical calculation software or data processing software. Given external environmental parameters, the pressure coefficient distribution data of the target spanwise position is calculated and exported. By comparing with the airfoil leading edge point, a set of coordinate points at different locations is obtained. Then, the surface pressure coefficients from different coordinate point sets are extracted. Finally, the surface pressure coefficients are interpolated sequentially to the chord position required for load calculation, obtaining the chord position surface pressure data of the high-camber wing. By rearranging the coordinate points and using different methods to distinguish the upper and lower surface points in different regions, the surface data of the wing (including high-camber wings) is successfully processed. The influence of wing camber variation is considered, which can meet the design requirements in engineering.

[0064] Finally, the following points should be noted: First, in the description of this application, it should be noted that, unless otherwise specified and limited, the terms "installation", "connection", and "linkage" should be interpreted broadly, and can be mechanical or electrical connections, or internal connections between two components, or direct connections. "Up", "down", "left", "right", etc. are only used to indicate relative positional relationships. When the absolute position of the described object changes, the relative positional relationship may change.

[0065] Secondly: The accompanying drawings of the embodiments disclosed in this invention only involve the structures involved in the embodiments disclosed in this invention. Other structures can refer to the general design. In the absence of conflict, the same embodiment and different embodiments of this invention can be combined with each other.

[0066] In conclusion, the above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the protection scope of the present invention.

Claims

1. An aerodynamic data processing method for a large-camber wing, characterized in that, include: Determine the target spanwise position for load design, extract the spanwise coordinate set of the target spanwise position based on the aircraft body coordinate system, and mark each coordinate point in the spanwise coordinate set; Input the wing target spanwise data into numerical calculation software or data processing software, give external environmental parameters, calculate and export the pressure coefficient distribution data of the target spanwise position; The coordinate points in the pressure coefficient distribution data are compared with the airfoil leading edge points in sequence according to the marking order to determine the airfoil leading edge points in the spanwise coordinate group. Then, the points that belong to the upper surface of the airfoil are formed into the first set, and the coordinate points that belong to the lower surface of the airfoil are formed into the second set. The surface pressure coefficients of the coordinate points in the first set and the second set are extracted. The surface pressure coefficients are interpolated sequentially to the chordal positions required for aerodynamic load calculation to obtain the chordal position data of the large-camber wing. The specific calculation method for the pressure coefficient distribution data is as follows: First, the pressure coefficient is calculated based on the external environmental parameters. Then, the three-dimensional coordinates and corresponding pressure coefficients at the spanwise profile are derived based on the target spanwise position. The basic form is (x,y,z,Cp). The specific methods for determining the data within the first and second sets are as follows: The pressure coefficient distribution data in the basic form (x, y, z, Cp) are sorted in ascending order according to the x-coordinate; The first point after sorting is taken as the leading edge coordinate point of the airfoil; Points located in the first certain proportion of the sorted array are determined to be at the leading edge position of the airfoil at that spanwise position. If their z-coordinate is greater than the z-coordinate of the first point, then the point belongs to the upper surface of the airfoil at that spanwise position and is included in the first set; otherwise, the point belongs to the lower surface of the airfoil at that spanwise position and is included in the second set. The remaining coordinate points are determined to be located in the mid-rear part of the airfoil in the spanwise direction. The first remaining coordinate point is selected according to the marked order, and its relative position to the last point on the upper surface that has already been determined is calculated: ； Among them, z up This is the last point on the upper surface that was previously determined to be complete, z down This is the last point on the lower surface that was previously determined to be complete; K0 is the target threshold, if K i If the value is less than K0, then point i is considered to be on the upper surface of the airfoil and included in the first set; otherwise, point i is determined to be on the lower surface of the airfoil and included in the second set. The position of the next coordinate point is determined according to the marked order until the remaining coordinate points have been determined.

2. The aerodynamic data processing method for a large-camber wing as described in claim 1, characterized in that: The spanwise profile of the target spanwise position is parallel to or at a certain angle to the xoz plane of the aircraft body coordinate system.

3. The aerodynamic data processing method for high-camber wings as described in claim 1, characterized in that, The specific method for obtaining the chordal position is as follows: Calculate the relative chord length of the corresponding positions of the coordinate points in the first and second sets using the chord length calculation formula. The chord length calculation formula is as follows: ； Among them, xc i Let x be the relative chord length at point i. i Let x be the x-coordinate of point i, and x1 be the x-coordinate of the leading edge point. m The x-coordinate of the trailing edge point; The relative distance is calculated using the relative chord length, and it is then determined whether the distance is less than a distance threshold. The determination formula is: ； If so, then the relative chord length is increased according to the distance threshold, as shown in the formula: ； If not, calculate the relative position of the corresponding coordinate point based on the load, perform linear interpolation on the arrays (xc, Cp) of the upper and lower surfaces respectively, and output the calculation results as required.

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