Aerodynamic coefficient mass ejection correction method and system based on wall surface friction resistance data

Abstract

The invention relates to the technical field of aircraft aerodynamic characteristic prediction, in particular to an aerodynamic coefficient mass ejection correction method and system based on wall surface friction resistance data, and the method comprises the steps: obtaining the fluid mechanics data of a target aircraft under the condition of no mass ejection, calculating the local streamline length s at each grid unit i of the wall surface based on the wall surface shear stress data; based on the wall surface pressure data, the shear stress data and the incoming flow parameters, calculating boundary layer outer edge parameters at each grid unit i, and based on the boundary layer outer edge parameters and the local streamline length s at each grid unit i, calculating the friction resistance of each grid unit i by referring to a friction resistance calculation model subjected to mass ejection correction. Calculating wall surface friction resistance data of each grid unit i under the condition of mass ejection; the wall surface friction data of all the grid units under the mass ejection condition are utilized, the total friction of the target aircraft is obtained through integration and is non-dimensionalized, and the aerodynamic coefficient is obtained. According to the invention, the prediction accuracy of the reference temperature method considering quality injection correction in a high-altitude and high-speed state is improved.

Description

Technical Field

[0001] This invention relates to the field of aircraft aerodynamic characteristic prediction technology, and in particular to a method and system for mass ejection correction of aerodynamic coefficients based on wall friction data. Background Technology

[0002] In aircraft design, obtaining aerodynamic data quickly and accurately is one of the primary tasks, as this data forms the basis for aerodynamic layout optimization, control system design, and flight simulation. Currently, the main methods for obtaining relatively accurate aerodynamic data are wind tunnel testing and computational fluid dynamics (CFD) simulation. However, both methods are generally time-consuming and costly, making it difficult to meet the urgent needs of large-scale data calculations and iterative analysis required in the preliminary and optimization design stages of aircraft. In contrast, aerodynamic engineering prediction methods, with their advantages of fast calculation speed and relatively accurate results, are widely used in the preliminary and optimization design of aircraft.

[0003] For aircraft flying at high altitudes and high speeds, frictional drag accounts for a large proportion of the total drag, and this proportion increases rapidly with increasing flight altitude, having a decisive impact on the lift-to-drag ratio. To effectively reduce frictional drag and aerodynamic heating, a common technique is "mass ejection," which involves actively injecting gas into the external flow field through the aircraft wall. Mass ejection significantly alters the state of the aircraft's boundary layer, thereby achieving drag reduction and heat dissipation. Furthermore, under extremely high-speed flight conditions, intense aerodynamic heating causes pyrolysis of thermal protection materials, also resulting in passive mass ejection. It is worth noting that in the initial design phase of an aircraft, aerodynamic predictions are typically calculated based on a "clean" shape without mass ejection, without considering the impact of mass ejection. If the mass ejection effect needs to be evaluated in the later stages of design, large-scale aerodynamic calculations must be performed again, which will significantly extend the design cycle and increase development costs considerably.

[0004] Among existing engineering prediction methods, the reference temperature method is the mainstream method for calculating the friction resistance of high-altitude, high-speed aircraft. This method uses the friction resistance calculation formula for incompressible boundary layers to calculate the friction resistance of compressible boundary layers, and the parameters in the formula are calculated using an artificially constructed boundary layer reference temperature. For flat plate flows, the reference temperature method can still give satisfactory results at low altitudes and in hyposonic conditions.

[0005] However, the traditional reference temperature method has significant limitations and cannot meet the requirements of the refined design of modern high-speed aircraft. Specifically, these limitations are as follows: First, the supersonic flight of blunt-nosed aircraft generates shock waves, and the boundary layer outer edge parameters differ significantly from the incoming flow parameters. If the reference temperature method directly uses the incoming flow parameters for calculation, it will lead to a large error. Second, the friction calculation formula of the reference temperature method requires the local Reynolds number of the current calculation point, and the selection of the length when calculating the Reynolds number affects the calculation accuracy. Third, high-altitude and high-speed flight brings a strong viscous interference effect, which affects the wall friction and pressure distribution. The reference temperature method does not consider the impact of this effect.

[0006] Therefore, there is an urgent need to develop a new method that can quickly and accurately correct and evaluate the mass ejection effect based on existing massless ejection aerodynamic data, in order to overcome the shortcomings of existing technologies, shorten the design cycle, and reduce R&D costs.

[0007] In view of this, the present invention is proposed. Summary of the Invention

[0008] The purpose of this invention is to provide a method and system for correcting aerodynamic coefficient mass ejection based on wall friction data, which solves the problem that the reference temperature method cannot be applied to situations with wall mass ejection, and improves the prediction accuracy of the reference temperature method under high-altitude and high-speed conditions.

[0009] In a first aspect, the present invention provides a method for mass ejection correction of aerodynamic coefficients based on wall friction data, comprising the following steps: S1. Acquire the hydrodynamic data of the target aircraft under massless ejection conditions, wherein the hydrodynamic data includes wall mesh data, wall shear stress data and wall pressure data; S2. Calculate each mesh element of the wall based on the wall shear stress data. i Local streamline length at the location s ; S3. Based on wall pressure data, shear stress data, and incoming flow parameters, calculate the values ​​for each mesh element. i The boundary layer outer edge parameters at the location, wherein the boundary layer outer edge parameters include the boundary layer outer edge pressure. Boundary layer outer edge density and boundary layer outer edge Mach number Ma e ; S4, Based on boundary layer outer edge parameters and individual mesh elements i Local streamline length at the location s Referring to the friction calculation model corrected by mass ejection, the friction of each grid cell is calculated. i Wall friction data under conditions of mass ejection; S5. Using the wall friction data of all grid cells under mass ejection conditions, the total friction of the target aircraft is obtained by integration and dimensionless transformation, thus obtaining the aerodynamic coefficients.

[0010] As a preferred embodiment of this technical solution, in step S2, each mesh unit of the wall surface i Local streamline length at the location s The calculation methods include: The limiting streamlines are calculated based on wall shear stress data, where the limiting streamlines satisfy the following equation:

[0011] In the formula, These are wall grid points. x、y、z Shear stress data in the direction; The above equations are solved using the fourth-order Runge-Kutta method to obtain the current element. i The streamlines to the stagnation point at the nose of the aircraft were calculated to obtain the mesh elements of each wall surface. i Local streamline length at the location s .

[0012] As a preferred embodiment of this technical solution, in step S3, each grid cell... i Boundary layer outer edge pressure With the corresponding wall pressure equal.

[0013] In a preferred embodiment of this technical solution, in step S3, the pressure after the shock wave is calculated based on the total pressure relationship before and after the shock wave. The pressure relationship before and after the shock wave is as follows:

[0014] In the formula, For the pressure of free flow, For the Mach number of the incoming flow, is the specific heat ratio of the gas.

[0015] In a preferred embodiment of this technical solution, in step S3, each grid cell is calculated based on the fact that the entropy along the streamline remains constant. i Boundary layer outer edge density The isentropic relationship is as follows:

[0016] In the formula, This represents the gas density after the shock wave.

[0017] As a preferred embodiment of this technical solution, in step S3, each grid cell... i Mach number at the outer edge of the boundary layer Mae The method for determining it is as follows: Set the Mach number at the outer edge of the boundary layer Ma e The initial assumed values; At the outer edge of the boundary layer, the parameter is Under these conditions, the wall friction coefficient was calculated using the reference temperature method with massless ejection. ; The current mesh element is calculated using wall shear stress data. i coefficient of friction at the point In the formula, ρ ∞ For the incoming flow density, u ∞ The incoming flow velocity; Continuously adjust the Mach number at the outer edge of the boundary layer Ma e Values, until and If the error between the two values ​​is less than a preset threshold, the Mach number used at this point is the final boundary layer outer edge Mach number. Ma e ; Preferably, the Mach number at the outer edge of the boundary layer is continuously adjusted. Ma e Values, until and When they are equal, the Mach number used at this point is the final boundary layer outer edge Mach number. Ma e .

[0018] As a preferred embodiment of this technical solution, step S4 includes: Based on boundary layer outer edge parameters The reference temperature is calculated using the following formula. T * :

[0019] In the formula, T w The wall temperature; based on p e and T * The density of the reference state is calculated using the following formula. :

[0020] In the formula, R is the gas constant; Density based on reference state Boundary layer outer edge velocity and grid cellsi Local streamline length at the location s The local Reynolds number is calculated using the following formula. Re s :

[0021] In the formula, μ This refers to the gas viscosity coefficient. Based on the friction calculation model corrected for mass ejection, the friction coefficient under mass ejection conditions is calculated. ; according to , to obtain mesh cells i The friction at that point has dimensions.

[0022] As a preferred embodiment of this technical solution, the friction calculation model after mass ejection correction in step S4 is as follows:

[0023] In the formula, , , Re s For the local Reynolds number, a n These are the polynomial coefficients.

[0024] In a preferred embodiment of this technical solution, in step S5, the frictional resistance is dimensionless based on the actual incoming flow pressure.

[0025] Secondly, the present invention also discloses an aerodynamic coefficient mass ejection correction system based on wall friction data, comprising: The data acquisition and processing module is used to acquire wall mesh data, wall shear stress data, and wall pressure data of the target aircraft under massless ejection conditions. The local streamline length calculation module is used to calculate the individual mesh elements of the wall based on the wall shear stress data. i Local streamline length at the location s ; The boundary layer outer edge parameter calculation module is used to calculate the parameters of each mesh cell based on wall pressure data and incoming flow parameters. i Boundary layer outer edge parameters at the location; The mass ejection friction calculation module is used to calculate friction based on boundary layer outer edge parameters and individual mesh elements. i Local streamline length at the location s Referring to the friction calculation model corrected by mass ejection, the friction of each grid cell is calculated. i Wall friction data under conditions of mass ejection; The aerodynamic coefficient generation module is used to integrate and dimensionlessly process the wall friction data of all grid cells under mass ejection conditions to generate aerodynamic coefficients of the aircraft that take into account the mass ejection effect.

[0026] Thirdly, the present invention also provides an electronic device, including a memory and a processor, wherein the memory is used to store one or more computer instructions, and the processor is used to execute one or more computer instructions for: Incoming flow conditions are Read in the wall mesh data and wall shear stress data of the target aircraft under massless ejection conditions. and wall pressure data ; Based on wall shear stress data, the limit streamline tracing method is used to calculate each mesh element of the wall. i streamline length at s ; Based on wall pressure data and inflow parameters, each grid cell is calculated using an iterative inversion method. i Boundary layer outer edge pressure Boundary layer outer edge density and boundary layer outer edge Mach number Ma e ; Based on the reconstructed boundary layer outer edge parameters and individual mesh elements i Local streamline length at the location s Combining the friction calculation model corrected by mass ejection, the friction of each grid cell is calculated. i Wall friction data under conditions of mass ejection; The wall friction data of all grid cells under mass ejection conditions are integrated and dimensionless to generate aerodynamic coefficients of the aircraft that take into account the mass ejection effect.

[0027] The aerodynamic coefficient mass ejection correction method based on wall friction data of the present invention has at least the following beneficial effects: This invention presents a method for correcting the mass ejection effect of aerodynamic coefficients based on wall friction data. Addressing the mass ejection effect caused by gas injection into the aircraft wall, it uses wall friction distribution data without gas injection and combines it with engineering methods to calculate aerodynamic data under the mass ejection effect. Based on existing wall friction and pressure distributions under no-gas-injection conditions, a boundary layer outer edge reference state parameter is constructed. Using the obtained boundary layer outer edge reference state parameter and wall mass flow rate, a reference temperature method with mass ejection correction is employed to calculate wall friction data under mass ejection conditions. First, a new reference state parameter is constructed using existing friction and pressure data under no-mass-ejection conditions. Calculating friction under mass ejection conditions using this new reference state parameter indirectly considers the influence of local parameters and viscous interference effects, improving the accuracy of friction prediction under mass ejection conditions compared to directly using the reference temperature state. Second, the constructed correction formula for friction calculation under mass ejection conditions based on dimensionless mass ejection quantity and Reynolds number broadens the application scope of the reference temperature method, enabling rapid prediction of friction under mass ejection conditions.

[0028] Therefore, this invention allows for the rapid assessment of the impact of mass ejection effects based on existing flow field data without recalculation, and the data is more accurate compared to the original reference temperature method. It overcomes the limitations of traditional engineering methods for calculating friction, which cannot consider viscous interference effects and local parameters, thus improving the accuracy of engineering prediction methods for mass ejection friction correction. Attached Figure Description

[0029] To more clearly illustrate the specific embodiments of the present invention or the technical solutions in the prior art, the drawings used in the description of the specific embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are some embodiments of the present invention. For those skilled in the art, other drawings can be obtained from these drawings without creative effort.

[0030] Figure 1 This is a schematic diagram of the mass ejector flow field of the present invention; Figure 2 This is a schematic diagram of the limiting streamline of the present invention. Detailed Implementation

[0031] The technical solution of the present invention will be clearly and completely described below with reference to the embodiments. 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.

[0032] In the description of this invention, it should be understood that the terms "center," "longitudinal," "lateral," "length," "width," "thickness," "upper," "lower," "front," "rear," "left," "right," "vertical," "horizontal," "top," "bottom," "inner," "outer," "clockwise," and "counterclockwise," etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are used only for the convenience of describing this invention and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limiting this invention.

[0033] Furthermore, the terms "first" and "second" are used for descriptive purposes only and should not be construed as indicating or implying relative importance or implicitly specifying the number of indicated technical features. Thus, a feature defined as "first" or "second" may explicitly or implicitly include one or more of the stated features. In the description of this invention, "a plurality of" means two or more, unless otherwise explicitly specified. Furthermore, the terms "installed," "connected," and "linked" should be interpreted broadly; for example, they may refer to a fixed connection, a detachable connection, or an integral connection; they may refer to a mechanical connection or an electrical connection; they may refer to a direct connection or an indirect connection through an intermediate medium; and they may refer to the internal connection of two components. Those skilled in the art can understand the specific meaning of the above terms in this invention based on the specific circumstances.

[0034] Example 1 The accuracy loss of friction prediction based on the reference temperature method for complex aircraft at high altitudes and high speeds mainly comes from three aspects: difficulty in accurately obtaining boundary layer outer edge parameters, length selection when calculating local Reynolds number, and viscous interference effect.

[0035] This embodiment proposes a mass ejection correction method for aerodynamic coefficients based on wall friction data. Firstly, since the wall velocity is 0 in the viscosity calculation, the wall streamline cannot be directly calculated. Therefore, by using the existing wall shear stress distribution and employing a streamline tracing method, the wall limiting streamline is obtained, with the local streamline length as the reference. s Calculate the local Reynolds number: Compared to directly calculating the Reynolds number using the distance from the aircraft's nose, calculating it using the local streamline length better reflects the fluid's journey from an undisturbed state to the current calculation point. Secondly, similar to the reference temperature method which calculates reference state parameters using an artificially constructed reference temperature, a new reference state is established based on existing wall friction data, pressure data, and incoming flow parameters. The parameters of this reference state are... Frictional resistance under mass ejection conditions is calculated using this reference state. Since the new reference state incorporates local pressure and frictional resistance, and to some extent considers the influence of viscous interference effects, it is more accurate than the classical reference temperature method. Finally, to establish an engineering prediction method for frictional resistance considering mass ejection, a dimensionless mass ejection quantity is used. With Reynolds number The polynomial function form modifies the friction calculation formula used in the classical reference temperature method, which is: The friction calculation model after mass ejection correction is as follows:

[0036] In the formula, , , Re s For the local Reynolds number, a n These are the polynomial coefficients.

[0037] like Figure 1-2 As shown, the aerodynamic coefficient mass ejection correction method based on wall friction data in this embodiment specifically includes the following steps: S1, under the condition of incoming flow: The study aims to acquire fluid dynamics data of the target aircraft under massless ejection conditions (viscous flow), including wall mesh data and wall shear stress data. and wall pressure data ; S2. Calculate each mesh element of the wall based on the wall shear stress data. i Local streamline length at the location s ; Specifically, each grid unit on the wall i Local streamline length at the location s The calculation methods include: First, since the wall velocity in viscous flow is zero, the limiting streamline represents the streamline of fluid particles approaching the wall in viscous flow. Therefore, the limiting streamline can be calculated using wall shear stress data. The limiting streamline satisfies the following equation:

[0038] In the formula, These are wall grid points. x、y、z Shear stress data in the direction; Then, the above equations are solved using the fourth-order Runge-Kutta method to obtain the current element. i The streamlines to the stagnation point at the nose of the aircraft are used to calculate the individual mesh elements of the wall surface. i Local streamline length at the locations S3. Based on wall pressure data and incoming flow parameters, calculate the values ​​for each grid cell. i The boundary layer outer edge parameters at the location, wherein the boundary layer outer edge parameters include the boundary layer outer edge pressure. Boundary layer outer edge density and boundary layer outer edge Mach number Ma e ; Specifically, according to boundary layer flow theory, the normal pressure gradient within the boundary layer is negligible, therefore each grid cell... i Boundary layer outer edge pressure With the corresponding wall pressure equal.

[0039] Specifically, the pressure after the shock wave is calculated based on the total pressure relationship before and after the shock wave. The pressure relationship before and after the shock wave is as follows: , In the formula, For the pressure of free flow, For the Mach number of the incoming flow, Specific heat ratio of gases; Furthermore, based on the assumption that the entropy along the streamline remains constant, the calculations are performed for each grid cell. i Boundary layer outer edge density The isentropic relationship is as follows:

[0040] In the formula, The density of the gas after the shock wave; Thus, the pressure at the outer edge of the boundary layer was obtained. With density .

[0041] Based on the above technical solution, further obtain each grid cell i Mach number at the outer edge of the boundary layer Ma e .

[0042] Specifically, first, the Mach number at the outer edge of the boundary layer is set. Ma e The initial assumption is that the Mach number at the outer edge of the boundary layer is equal to the incoming Mach number; At the outer edge of the boundary layer, the parameter is Under these conditions, the wall friction coefficient was calculated using the reference temperature method with massless ejection. ; The current mesh element is calculated using wall shear stress data. i coefficient of friction at the point In the formula, ρ ∞ For the incoming flow density, u ∞ For the incoming flow velocity, These are wall grid points. x、y、z Shear stress data in the direction; Continuously adjust the Mach number at the outer edge of the boundary layer Ma e Values, until and If the error between the two values ​​is less than a preset threshold, the Mach number used at this point is the final boundary layer outer edge Mach number. Ma e ; More preferably, the Mach number at the outer edge of the boundary layer is continuously adjusted. Ma e Values, until and When they are equal, the Mach number used at this point is the final boundary layer outer edge Mach number. Ma e .

[0043] S4, Based on boundary layer outer edge parameters and individual mesh elements i Local streamline length at the location s Referring to the friction calculation model corrected by mass ejection, the friction of each grid cell is calculated. i Wall friction data under conditions of mass ejection; Specifically, firstly, based on the boundary layer outer edge parameters The reference temperature is calculated using the following formula. T * :

[0044] In the formula, T w The wall temperature is known. in, T e It can be derived from the gas law. p e =ρ e RT e We obtain the equation, where R is the gas constant, and for air R = 287.06 J / (kg˙K).

[0045] Furthermore, based on p e and T * The density of the reference state is calculated using the following formula. :

[0046] In the formula, R is the gas constant, and for air R = 287.06 J / (kg˙K); Furthermore, density based on the reference state Boundary layer outer edge velocity and grid cells i Local streamline length at the location s The local Reynolds number is calculated using the following formula. Re s :

[0047] In the formula, μ This refers to the gas viscosity coefficient. Among them, the outer edge velocity of the boundary layer can be Calculated.

[0048] Then, based on the friction calculation model corrected for mass ejection, the friction coefficient under the condition of mass ejection is calculated. ; The friction calculation model after mass ejection correction is as follows:

[0049] In the formula, , , Re s For the local Reynolds number, a n These are the polynomial coefficients.

[0050] Finally, according to Calculate the mesh cells i The friction at that point has dimensions.

[0051] S5. Following steps S1-S5, obtain wall friction data for all grid cells under mass ejection conditions. Using all grid cells under mass ejection conditions, integrate to obtain the total friction of the target aircraft. Finally, use the actual incoming flow pressure to make the friction dimensionless, generating the aerodynamic coefficient of the aircraft considering the mass ejection effect.

[0052] Based on benchmark CFD data, this invention can quickly and reliably predict the drag reduction effect of various mass ejection schemes, greatly accelerating the design and optimization process of aircraft.

[0053] Example 2 This embodiment provides an aerodynamic coefficient mass ejection correction system based on wall friction data, including: The data acquisition and processing module is used to acquire wall mesh data, wall shear stress data, and wall pressure data of the target aircraft under massless ejection conditions. The local streamline length calculation module is used to calculate the individual mesh elements of the wall based on the wall shear stress data. i Local streamline length at the location s ; The boundary layer outer edge parameter calculation module is used to calculate the parameters of each grid cell based on conditions such as wall friction, pressure, and incoming flow parameters without mass ejection, using aerodynamic principles and tools such as the reference temperature method and isentropic conditions. i Boundary layer outer edge parameters at the location; The mass ejection friction calculation module is used to calculate friction based on boundary layer outer edge parameters and individual mesh elements. i Local streamline length at the location s Referring to the friction calculation model corrected by mass ejection, the friction of each grid cell is calculated. i Wall friction data under conditions of mass ejection; The aerodynamic coefficient generation module is used to integrate and dimensionlessly process the wall friction data of all grid cells under mass ejection conditions to generate aerodynamic coefficients of the aircraft that take into account the mass ejection effect.

[0054] Example 3 This embodiment provides an electronic device, including a memory and a processor, wherein the memory is used to store one or more computer instructions, and the processor is used to execute one or more computer instructions for: Incoming flow conditions are Read in the wall mesh data and wall shear stress data of the target aircraft under massless ejection conditions. and wall pressure data ; Based on wall shear stress data, the limit streamline tracing method is used to calculate each mesh element of the wall. i Local streamline length at the location s ; Based on wall pressure data and inflow parameters, each grid cell is calculated using an iterative inversion method. i Boundary layer outer edge pressure Boundary layer outer edge density and boundary layer outer edge Mach number Ma e ; Based on the reconstructed boundary layer outer edge parameters and individual mesh elements i Local streamline length at the location s Combining the friction calculation model corrected by mass ejection, the friction of each grid cell is calculated. i Wall friction data under conditions of mass ejection; The wall friction data of all grid cells under mass ejection conditions are integrated and dimensionless to generate aerodynamic coefficients of the aircraft that take into account the mass ejection effect.

[0055] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention, and not to limit them; although the present invention has been described in detail with reference to the foregoing embodiments, those skilled in the art should understand that modifications can still be made to the technical solutions described in the foregoing embodiments, or equivalent substitutions can be made to some or all of the technical features; and these modifications or substitutions do not cause the essence of the corresponding technical solutions to deviate from the scope of the technical solutions of the embodiments of the present invention.

Claims

1. A method for mass ejection correction of aerodynamic coefficients based on wall friction data, characterized in that, Includes the following steps: S1. Acquire the hydrodynamic data of the target aircraft under massless ejection conditions, wherein the hydrodynamic data includes wall mesh data, wall shear stress data and wall pressure data; S2. Calculate each mesh element of the wall based on the wall shear stress data. i Local streamline length at the location s ; S3. Based on wall pressure data, shear stress data, and incoming flow parameters, calculate the values ​​for each mesh element. i The boundary layer outer edge parameters at the location, wherein the boundary layer outer edge parameters include the boundary layer outer edge pressure. Boundary layer outer edge density and boundary layer outer edge Mach number Ma e ; S4, Based on boundary layer outer edge parameters and individual mesh elements i Local streamline length at the location s Referring to the friction calculation model corrected by mass ejection, the friction of each grid cell is calculated. i Wall friction data under conditions of mass ejection; S5. Using the wall friction data of all grid cells under mass ejection conditions, the total friction of the target aircraft is obtained by integration and dimensionless transformation, thus obtaining the aerodynamic coefficients.

2. The aerodynamic coefficient mass ejection correction method based on wall friction data according to claim 1, characterized in that, In step S2, each mesh element of the wall surface i Local streamline length at the location s The calculation methods include: The limiting streamlines are calculated based on wall shear stress data, where the limiting streamlines satisfy the following equation: In the formula, These are wall grid points. x, y, and z Shear stress data in the direction; The above equations are solved using the fourth-order Runge-Kutta method to obtain the current element. i The streamlines to the stagnation point at the nose of the aircraft were calculated to obtain the mesh elements of each wall surface. i Local streamline length at the location s .

3. The aerodynamic coefficient mass ejection correction method based on wall friction data according to claim 1, characterized in that, In step S3, each grid cell i Boundary layer outer edge pressure With the corresponding wall pressure equal.

4. The aerodynamic coefficient mass ejection correction method based on wall friction data according to claim 1, characterized in that, In step S3, the pressure after the shock wave is calculated based on the total pressure relationship before and after the shock wave. The pressure relationship before and after the shock wave is as follows: In the formula, For the pressure of free flow, For the Mach number of the incoming flow, is the specific heat ratio of the gas.

5. The aerodynamic coefficient mass ejection correction method based on wall friction data according to claim 1, characterized in that, In step S3, based on the assumption that the entropy along the streamline remains constant, the entropy of each grid cell is calculated. i Boundary layer outer edge density The isentropic relationship is as follows: In the formula, This represents the gas density after the shock wave.

6. The aerodynamic coefficient mass ejection correction method based on wall friction data according to claim 1, characterized in that, In step S3, each grid cell i Mach number at the outer edge of the boundary layer Ma e The method for determining it is as follows: Set the Mach number at the outer edge of the boundary layer Ma e The initial assumed values; At the outer edge of the boundary layer, the parameter is Under these conditions, the wall friction coefficient was calculated using the reference temperature method with massless ejection. ; The current mesh element is calculated using wall shear stress data. i coefficient of friction at the point In the formula, ρ ∞ For the incoming flow density, u ∞ The incoming flow velocity; Continuously adjust the Mach number at the outer edge of the boundary layer Ma e Values, until and If the error between the two values ​​is less than a preset threshold, the Mach number used at this point is the final boundary layer outer edge Mach number. Ma e ; Preferably, the Mach number at the outer edge of the boundary layer is continuously adjusted. Ma e Values, until and When they are equal, the Mach number used at this point is the final boundary layer outer edge Mach number. Ma e .

7. The aerodynamic coefficient mass ejection correction method based on wall friction data according to claim 1, characterized in that, Step S4 includes: Based on boundary layer outer edge parameters The reference temperature is calculated using the following formula. T * : In the formula, T w The wall temperature; based on p e and T * The density of the reference state is calculated using the following formula. : In the formula, R is the gas constant; Density based on reference state Boundary layer outer edge velocity and grid cells i Local streamline length at the location s The local Reynolds number is calculated using the following formula. Re s : In the formula, μ This refers to the gas viscosity coefficient; Based on the friction calculation model corrected for mass ejection, the friction coefficient under mass ejection conditions is calculated. ; according to , to obtain mesh cells i The friction at that point is dimensional.

8. The aerodynamic coefficient mass ejection correction method based on wall friction data according to claim 1, characterized in that, In step S4, the friction calculation model after mass ejection correction is as follows: In the formula, , , Re s For the local Reynolds number, a n These are the polynomial coefficients.

9. The aerodynamic coefficient mass ejection correction method based on wall friction data according to claim 1, characterized in that, In step S5, the frictional resistance is dimensionless using the actual incoming flow pressure.

10. A mass ejection correction system for aerodynamic coefficients based on wall friction data, characterized in that, include: The data acquisition and processing module is used to acquire wall mesh data, wall shear stress data, and wall pressure data of the target aircraft under massless ejection conditions. The local streamline length calculation module is used to calculate the individual mesh elements of the wall based on the wall shear stress data. i Local streamline length at the location s ; The boundary layer outer edge parameter calculation module is used to calculate the parameters of each mesh element based on wall pressure data, shear stress data, and incoming flow parameters. i Boundary layer outer edge parameters at the location; The mass ejection friction calculation module is used to calculate friction based on boundary layer outer edge parameters and individual mesh elements. i Local streamline length at the location s Referring to the friction calculation model corrected by mass ejection, the friction of each grid cell is calculated. i Wall friction data under conditions of mass ejection; The aerodynamic coefficient generation module is used to integrate and dimensionlessly process the wall friction data of all grid cells under mass ejection conditions to generate aerodynamic coefficients of the aircraft that take into account the mass ejection effect.

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