A method and device for determining a cold gas point source of a turbine blade of an aero-engine, a computer device and a medium

Abstract

Embodiments of the present application provide a kind of turbine blade cold gas point source determination method, device, computer equipment and medium of aero-engine, it is related to turbine blade cooling design technical field, wherein, the method comprises the following steps: setting the attribute parameter of turbine blade, and attribute parameter is saved to input configuration file;The geometry analysis of the blade shape format of turbine blade is carried out, the positioning dimension of turbine blade is divided into blade surface feature position and blade upper and lower end area feature position, the three-dimensional coordinates of blade cold gas point source are determined according to positioning dimension;Set the cold gas injection angle corresponding to each blade cold gas point source, and save to input configuration file;The cold gas injection flow and cold gas injection temperature of each blade cold gas point source are set respectively, and save to input configuration file respectively.The accurate positioning of turbine blade surface and upper and lower end area wall surface cold gas point source, direction setting and parameter automatic loading are realized.

Description

Technical Field

[0001] This invention relates to the field of turbine blade cooling design technology, and in particular to a method, apparatus, computer equipment, and medium for determining the cooling gas point source of aero-engine turbine blades. Background Technology

[0002] In aero-engine turbine blade design, the geometric location, injection direction, temperature, and flow rate parameters of cool gas point sources (such as film cooling holes and cooling injection holes) are key factors affecting blade cooling efficiency and overall engine thermal efficiency. Traditional turbine blade cool gas point source loading relies on manual operation using 3D modeling tools such as UG and CATIA. The specific process involves designers manually extracting the geometric coordinates of the blade surface / end zone wall, defining the cool gas injection direction point by point, manually calculating the cool gas flow rate and temperature parameters, and finally generating a cooling boundary condition file. However, this traditional method has significant technical drawbacks:

[0003] 1. Low efficiency: The single-bladed exhaust cooling point source loading requires manual operation for more than 5 hours, and the design cycle of multi-bladed exhaust (such as 3-5 stage turbines) is as long as 20-25 hours, which cannot meet the design iteration requirements of aero-engines;

[0004] 2. Insufficient accuracy: Geometric coordinate extraction relies on the designer's experience, and the positional error can reach ±0.5mm. Due to the lack of a unified coordinate system transformation logic, the spray direction is prone to non-physical errors such as "reverse spraying into the inner channel," with an error rate exceeding 30%.

[0005] 3. Poor adaptability: It does not support automatic parsing of mainstream blade model formats such as COORD and Geomturbo. For special locations such as the leading edge, trailing edge, and end plate of the blade, the extraction logic needs to be redeveloped. It is not compatible with different types of turbine blades (stator / rotor, high pressure / low pressure).

[0006] 4. Iteration difficulties: When blade geometry parameters (such as chord length and spanwise height) or cooling schemes (such as the number of cooling gas points and injection angle) are modified, the entire process must be manually re-executed, resulting in extremely low iteration efficiency.

[0007] As the inlet temperature of aero-engine turbines continues to rise (reaching 1800-2300K), the requirements for the precision of blade cooling design are becoming increasingly stringent, and traditional manual loading methods can no longer meet engineering needs. Therefore, there is an urgent need for a technical method that can achieve parameterized automatic loading of cold gas point sources, high-precision matching, and multi-scenario adaptability, to solve the pain points of low efficiency, poor accuracy, and difficulty in iteration of traditional methods. Summary of the Invention

[0008] In view of this, embodiments of the present invention provide a method for determining the cold gas point source of aero-engine turbine blades, to solve the technical problems of low efficiency, poor accuracy, weak adaptability, and difficulty in iteration in existing methods for loading cold gas point sources for aero-engine turbine blades. The method includes:

[0009] Initialize the input configuration file, set the attribute parameters of the turbine blades, and save the attribute parameters to the input configuration file;

[0010] Geometric analysis is performed on the blade profile of the turbine blade, and the positioning dimension of the turbine blade is divided into the blade surface feature position and the blade upper and lower end area feature position. Based on the blade geometry file, the three-dimensional coordinates of the blade cooling gas point source are determined according to the positioning dimension. The blade cooling gas point source includes the blade surface cooling gas point source and the blade upper and lower end area cooling gas point source.

[0011] Set the air injection angle corresponding to each of the blade air source points, and save the three-dimensional coordinates of the blade air source points and the air injection angle to the input configuration file. The air injection angle includes the circumferential angle. α With radial angle β ;

[0012] Set the air injection flow rate and air injection temperature for each of the blade air source points, and save the air injection flow rate and air injection temperature to the input configuration file respectively.

[0013] This invention also provides a device for determining the cold gas point source of aero-engine turbine blades, to solve the technical problems of low efficiency, poor accuracy, weak adaptability, and difficulty in iteration in existing methods for loading cold gas point sources for aero-engine turbine blades. The device includes:

[0014] The turbine blade parameter setting module is used to initialize the input configuration file, set the attribute parameters of the turbine blade, and save the attribute parameters to the input configuration file.

[0015] The cooling gas point source geometric position setting module is used to perform geometric analysis on the blade shape format of the turbine blade, divide the positioning dimension of the turbine blade into the blade surface feature position and the blade upper and lower end area feature position, and determine the three-dimensional coordinates of the blade cooling gas point source based on the blade geometry file and the positioning dimension.

[0016] The injection direction setting module is used to set the cold air injection angle corresponding to each of the blade cold air point sources, and save the three-dimensional coordinates of the blade cold air point sources and the cold air injection angle to the input configuration file, wherein the cold air injection angle includes the circumferential angle. α With radial angle β ;

[0017] The cold air injection parameter setting module is used to set the cold air injection flow rate and cold air injection temperature for each of the cold air point sources on the blades, and save the cold air injection flow rate and the cold air injection temperature to the input configuration file respectively.

[0018] This invention also provides a computer device, including a memory, a processor, and a computer program stored in the memory and executable on the processor. When the processor executes the computer program, it implements the above-mentioned method for determining the cold gas point source of any aero-engine turbine blade, thereby solving the technical problems of low efficiency, poor accuracy, weak adaptability, and difficulty in iteration in the existing aero-engine turbine blade cold gas point source loading method.

[0019] This invention also provides a computer-readable storage medium storing a computer program that executes any of the above-described methods for determining the cold gas point source of aero-engine turbine blades, in order to solve the technical problems of low efficiency, poor accuracy, weak adaptability, and difficulty in iteration in the prior art of loading cold gas point sources for aero-engine turbine blades.

[0020] Compared with the prior art, the beneficial effects that at least one technical solution adopted in the embodiments of this specification can achieve include at least:

[0021] It achieves precise positioning, direction setting, and automatic parameter loading of the cold air point source on the surface of the turbine blades and the upper and lower end walls. Attached Figure Description

[0022] To more clearly illustrate the technical solutions of the embodiments of this application, the drawings used in the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of this application. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0023] Figure 1 This is a flowchart of a method for determining the cold gas point source of an aero-engine turbine blade, provided in an embodiment of the present invention.

[0024] Figure 2 This is a schematic diagram of the location of features on the blade surface provided in an embodiment of the present invention;

[0025] Figure 3 This is a schematic diagram of the characteristic positions of the upper and lower end regions of the blade provided in an embodiment of the present invention;

[0026] Figure 4 This is a structural block diagram of a computer device provided in an embodiment of the present invention;

[0027] Figure 5This is a structural block diagram of a device for determining the cold gas point source of an aero-engine turbine blade, provided in an embodiment of the present invention. Detailed Implementation

[0028] The embodiments of this application will now be described in detail with reference to the accompanying drawings.

[0029] The following specific examples illustrate the implementation of this application. Those skilled in the art can easily understand other advantages and effects of this application from the content disclosed in this specification. Obviously, the described embodiments are only a part of the embodiments of this application, and not all of them. This application can also be implemented or applied through other different specific embodiments, and the details in this specification can also be modified or changed based on different viewpoints and applications without departing from the spirit of this application. It should be noted that, in the absence of conflict, the following embodiments and features in the embodiments can be combined with each other. Based on the embodiments in this application, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of this application.

[0030] In this embodiment of the invention, a method for determining the cold gas point source of an aero-engine turbine blade is provided, such as... Figure 1 As shown, the method includes:

[0031] Step S101: Initialize the input configuration file, set the attribute parameters of the turbine blades, and save the attribute parameters to the input configuration file;

[0032] Step S102: Perform geometric analysis on the blade shape format of the turbine blade, divide the positioning dimension of the turbine blade into the blade surface feature position and the blade upper and lower end area feature position. Based on the blade geometry file, determine the three-dimensional coordinates of the blade cooling point source according to the positioning dimension, wherein the blade cooling point source includes the blade surface cooling point source and the blade upper and lower end area cooling point source.

[0033] Step S103: Set the air injection angle corresponding to each blade air source, and save the three-dimensional coordinates of the blade air source and the air injection angle to the input configuration file, wherein the air injection angle includes the circumferential angle. α With radial angle β ;

[0034] Step S104: Set the cold air injection flow rate and cold air injection temperature for each of the blade cold air point sources, and save the cold air injection flow rate and cold air injection temperature to the input configuration file respectively.

[0035] In specific implementation, the following steps are used to divide the positioning dimension of the turbine blade into the feature positions on the blade surface and the feature positions at the upper and lower ends of the blade. Based on the blade geometry file, the three-dimensional coordinates of the blade's cooling gas point source are determined according to the positioning dimension:

[0036] Based on the blade geometry file, the geometric entities of the turbine blade are identified, including the leading edge, trailing edge, suction surface, pressure surface, hub surface, casing surface, and flange connecting the blade to the endwall. Based on the cooling requirements of the turbine blade, the blade feature positions are determined, including blade surface feature positions and upper and lower end area feature positions. The blade surface feature positions and the upper and lower end area feature positions are mapped to the corresponding geometric entities in the blade geometry file. Parametric positioning logic is defined for the mapped blade surface feature positions and the upper and lower end area feature positions, respectively. The positioning logic for the blade surface feature positions includes dimensionless arc length and spanwise distribution parameters, and the positioning logic for the upper and lower end area feature positions includes end area anchor point coordinates and phase angle ranges. The three-dimensional coordinates of multiple blade surface cold air point sources are calculated using the dimensionless arc length and the spanwise distribution parameters. The three-dimensional coordinates of multiple upper and lower end area cold air point sources are also calculated using the end area anchor point coordinates and the phase angle ranges.

[0037] In specific implementation, the cooling requirements based on turbine blades are achieved through the following steps, including determining the characteristic locations of the blades:

[0038] Typical regions of thermal load on the turbine blade surface are identified, including the leading edge region, the pressure-side blade head region, and the trailing edge region. Based on these typical regions, the turbine blade surface is divided into multiple blade surface characteristic locations, including the suction-side leading edge, the pressure-side leading edge, the pressure-side blade head, and the trailing edge. The end-area cooling region and the tip clearance heat transfer region requiring cooling are identified in the blade-hub connection area. Based on these end-area cooling regions and tip clearance heat transfer regions, the upper and lower end areas of the blade are divided into multiple upper and lower end-area characteristic locations, including the hub blade inter-blade edge plate, the casing blade inter-blade edge plate, the hub blade leading edge plate, the casing blade leading edge plate, the hub blade trailing edge plate, the casing blade trailing edge plate, and the moving blade tip region.

[0039] In practice, the following steps are used to set the cold air injection angle corresponding to each blade cold air point source:

[0040] If the blade cooling air source is a cooling air source on the blade surface, the set radial angle β and circumferential angle αAs the cold air injection angle; if the cold air source of the blade is a cold air source in the upper or lower end region of the blade, the radial angle of the cold air injection angle is determined based on the end region position of the cold air source in the upper or lower end region of the blade. β Make corrections, and adjust the radial angle accordingly. β and the set circumferential angle α As the cold air injection angle, the end region position includes the upper end of the blade and the lower end of the blade.

[0041] In specific implementation, the radial angle of the cold air injection angle is determined by the following steps based on the end region positions of the cold air point sources located at the upper and lower end regions of the blade. β Make corrections:

[0042] If the cold air source in the upper and lower end regions of the blade is located at the lower end of the blade, the radial angle will be... β Set to positive; if the cold air source in the upper and lower end regions of the blade is located at the upper end of the blade, set the radial angle to positive. β Set to negative.

[0043] In specific implementation, the attribute parameters are achieved through the following steps:

[0044] Number of blade rows, blade type, blade data unit, and number of blade channels.

[0045] In specific implementation, the adaptive adjustment of the cold air injection flow rate and the cold air injection angle based on the engine operating status is achieved through the following steps:

[0046] The input configuration file defines multiple engine operating states and corresponding engine operating condition parameter sets for each engine operating state. The engine operating states include takeoff, cruise, and maneuvering states. Based on the current engine operating state and the corresponding engine operating condition parameter set, the blade thermal boundary condition parameter set and mainstream flow field information for the current engine operating state are determined from a preset association database. The blade thermal boundary condition parameter set includes the heat load reference distribution at characteristic locations on the blade surface and at the upper and lower end regions of the blade. The blade thermal boundary condition parameter set for the current engine operating state is compared with the blade thermal boundary condition parameter set for the preset reference state. The relative changes in cooling requirements at characteristic locations on the blade surface and at the upper and lower end regions of the blade are calculated, and the cold air injection flow rate of each blade cold air point source in the input configuration file is adjusted based on these relative changes. Based on the mainstream flow field information for the current engine operating state, the circumferential angle of the cold air injection angle of the cold air point sources on the blade surface is adjusted. α With radial angle β Make corrections; update the adjusted air injection flow rate and corrected air injection angle to the input configuration file.

[0047] In one embodiment of the present invention, the method for parameterizing the cooling gas point source of aero-engine turbine blades includes:

[0048] Step 1: Parametric analysis of turbine blade geometric model.

[0049] By configuring the Input.dat file (input configuration file), you can set the number of blade rows (e.g., 6 rows for a 3-stage high-pressure turbine), blade type (e.g., 1 stator blade and -1 rotor blade), blade data unit (e.g., millimeters / meters), and number of blade channels (set according to the number of blades and periodic characteristics, e.g., 36 blades correspond to 36 channels). Based on the provided general turbine design blade profile format, geometric analysis is performed, enabling automatic recognition of COORD format (including the flow channel file MERCOO) and Geomturbo format. Preset format verification rules (e.g., Geomturbo files have no extra blank lines, COORD files contain key blade profile coordinate fields) are used to exclude models with incorrect formats. By placing the blade geometry file and the "Input.dat" file in the same working directory and starting the analysis program, key blade profile coordinates (e.g., leading edge, trailing edge, blade head / blade feature points) are automatically extracted. If analysis errors occur (e.g., missing file, abnormal format), error logs are output in real time (e.g., "COORD file is missing MERCOO flow channel data") to guide the correction of the turbine blade geometry model.

[0050] Prepare the turbine blade geometry model that requires loading the cold air point source parameters. This loading program supports two mainstream blade geometry model formats: COORD, a common styling format for aero-engine turbine blade design, and Geomturbo. The COORD format requires the inclusion of the flow channel file MERCOO, while the Geomturbo format file needs to be in a standard format, version GEOMETRYTURBOVERSION 3.7. Locate the corresponding keyword fields in the Input.dat file and set parameters such as the number of blade rows, blade type, and blade data units. Specifically, first set the blade type: set the stator blade type to 1 and the rotor blade type to -1. Also, considering that there may be multiple blades in the same row, finally set the number of blade channels for that row to complete the parameterization and analysis of the geometry model.

[0051] Step 2: Parameterize the geometric position of the air conditioning point source.

[0052] Based on the geometric features of the blade, a two-dimensional parametric positioning system of "surface feature position + end region feature position" is constructed.

[0053] In the input configuration file, you can see information such as the type (blade surface / end zone wall), flow direction parameters, spanwise parameters, and cold air code for different locations. By selecting the type and inputting parameter values, you can complete the parameterization setting of the cold air point source location. The loading process of the cold air point source location on the blade surface and the point source location on the blade end zone will be introduced next.

[0054] In the dimension of locating the cold air source on the blade surface, four characteristic position codes are defined: 1# (suction-side leading edge), 2# (pressure-side leading edge), 3# (pressure-side blade base), and 4# (tail edge). A method using dimensionless parameters to define positions is proposed to improve adaptability to different blade shapes. With the leading edge as 0% and the tail edge as 100%, a dimensionless value along the blade arc length is input (e.g., 0.05 represents the 5% arc length position at the leading edge). The spanwise positions are automatically evenly distributed by setting the "number of spanwise points" (e.g., 15 points). For example, setting "dimensional arc length 0.05, number of spanwise points 15" for position 1# automatically calculates the three-dimensional coordinates (X, Y, Z) of 15 cold air points. In the dimension of locating the cold air source in the upper and lower end areas of the blade, seven characteristic position codes are defined: 5# (hub blade inter-edge plate), 6# (hub blade leading edge plate), 7# (casing blade inter-edge plate), 8# (casing blade leading edge plate), 9# (moving blade tip area), 10# (hub blade trailing edge plate), and 11# (casing blade trailing edge plate). Based on the periodic symmetry characteristics of the blade, the number of cold air points within the specified angle range can be calculated using the "end area anchor point coordinates (X,Y,Z)" and the "phase angle range (e.g., 10°)" (e.g., based on end area anchor point coordinates (0,0,300), 30 points, and a phase angle of 10°). The flow channel boundary is automatically identified based on the end area type (hub / casing) to achieve intelligent boundary matching, preventing cold air points from exceeding the flow channel range.

[0055] 2.1 Loading of cold air point source on blade surface.

[0056] Based on the flow direction along the blade's arc length from the leading edge to the trailing edge, four position codes were assigned to correspond to four different locations on the blade surface: position 1 corresponds to the area near the leading edge on the suction side of the blade, position 2 corresponds to the area near the leading edge on the pressure side of the blade, position 3 corresponds to the area near the blade head on the pressure side of the blade, and position 4 corresponds to the area near the trailing edge of the blade. Figure 2 As shown.

[0057] Based on the four defined characteristic locations, the geometric location of the cooling air point source is defined as a dimensionless parameter on the pressure and suction sides of the blade: with the leading edge of the blade as 0% and the trailing edge as 100%, the remaining locations on the pressure and suction sides are defined as dimensionless values ​​from 0% to 100% according to the design requirements and the proportion of the arc length of each side. Users can adjust this dimensionless parameter to determine the specific location according to the blade cooling efficiency design requirements. The specific location of this parameter is the first parameter of each characteristic location.

[0058] In the spanwise direction of the blade, the number of spanwise distribution points is given according to the blade's cooling efficiency design target. The loading of cooling gas point sources at different positions along the blade spanwise is automatically achieved through the point number parameter. By coordinating the setting of flow direction and spanwise parameters, the precise positioning of the turbine blade cooling gas point source at any geometric position along the blade surface can be quickly completed.

[0059] 2.2 Loading of cold air point source positions on the upper and lower end walls of turbine blades.

[0060] Along the axial direction, the characteristic position codes of the cold air on the upper and lower endwalls of the blade from the inlet to the outlet are defined. Seven position codes are set to correspond to seven different positions on the blade surface: position 5 corresponds to the wall position between blade rows on the lower endwall of the blade row; position 6 corresponds to the cold air position at the leading edge sealing the lower endwall of the blade row; position 7 corresponds to the wall position between blade rows on the upper endwall of the blade row; position 8 corresponds to the cold air position at the leading edge of the upper endwall of the blade row; position 9 corresponds to the cold air position in the tip region of the moving blade; position 10 corresponds to the cold air position at the trailing edge sealing the lower endwall of the blade row; and position 11 corresponds to the cold air position at the trailing edge sealing the upper endwall of the blade row. Figure 3 As shown.

[0061] Based on the defined seven feature locations, the geometric location of the cooling gas source is defined as the axial position parameter of the blade. The coordinates (X, Y, Z) of the corresponding feature locations are extracted from the vicinity of the periodic boundaries of the upper and lower flow channels of the blade as anchor points. Since the upper and lower endwalls have periodic symmetry features, it is only necessary to set the phase angle covered by the cooling gas points on the endwalls and the number of cooling gas points in that angle region. The system can automatically identify the geometric boundaries and positional features of the upper and lower end regions based on the feature location codes, accurately match the user-defined cooling gas point locations, and complete the loading without manual intervention in geometric boundary judgment.

[0062] Step 3: Set the direction of the cold air point source injection.

[0063] After step 2, the direction of the cold gas point source injection can be set. Similarly, this also needs to be done through the "Input.dat" file. Based on the distribution characteristics of the cold gas injection direction of aero-engine turbine blades and the design parameterized loading logic, the definition and automatic loading of the cold gas point source injection direction on the blade surface and the upper and lower end wall areas are implemented respectively. At this point, it is necessary to find and set the key fields for the geometric position of the cold gas point source.

[0064] Based on the geometric position of the cold air point source on the blade surface set in step 2, the direction of cold air injection at that point is set. From the perspective of the global coordinate system, the injection direction of the cold air point source on the blade surface varies along the blade arc length (from the leading edge to the trailing edge). However, the user-defined input airflow angle is the local coordinate-defined injection angle of the cold air point source at that point (including the circumferential angle α and the radial angle β). The same injection angle at different local coordinate systems may differ in the global coordinate system. Therefore, this program can automatically convert the angle parameters in the local coordinate system into the injection direction vector in the global coordinate system using a built-in coordinate system transformation algorithm, decomposing it into components in the X, Y, and Z directions. This ultimately achieves precise loading of the injection direction of the cold air point source at any arc length position on the blade surface, and the injection direction at different relative arc length positions can be independently customized to meet diverse design needs.

[0065] The loading method for the airflow injection direction of the cold air point source on the upper and lower end areas of the blade is similar to that of the cold air injection direction on the blade surface. The only difference is that the cold air injection direction in the upper and lower end areas is defined as an angle in the global coordinate system, thus eliminating the need for a transformation from local to global coordinates. Based on seven defined characteristic positions in the upper and lower end areas of the blade, the positions for the cold air injection direction parameters are defined, including the circumferential angle α and the radial angle β. According to the difference in airflow angle direction between the upper and lower end areas of the turbine blade, an automatic direction vector matching logic is built-in: for the cold air point on the lower end area wall, when the airflow is injected into the inner channel, the radial angle is set to positive, and the program outputs a positive radial component of the cold air injection direction vector; for the cold air point on the upper end area wall, when the airflow is injected into the inner channel, the radial angle is set to negative, and a negative radial component of the cold air injection direction vector is output. Simultaneously, cold air point sources are defined at any position along the blade inlet and outlet axial direction, enabling independent setting of the injection direction at different axial positions and ensuring the adaptability of the cold air injection direction in the upper and lower end areas to the airflow characteristics of the channel. Through parameterization, the injection angle corresponding to the axial position can be defined. Based on the differences between the upper and lower end regions (the radial component of the lower end region is positive by default, and that of the upper end region is negative), intelligent suggested values ​​are provided to complete the loading of the airflow direction of the cold air source on the end region blades.

[0066] Step 4: Parameterized calculation of cold air injection temperature and flow rate.

[0067] In addition to geometric location and orientation, turbine coolant point source parameters also need to include parameterized definitions of coolant injection temperature and flow rate. The loading of coolant temperature and flow rate has a significant impact on the accuracy of turbine performance evaluation. Due to the large number of coolant point source parameters for air-cooled turbines, a method for loading coolant flow rate as a relative percentage and for automatically selecting coolant temperature is proposed. By setting the number of key coolant locations, the automatic loading of coolant flow rate and temperature for each coolant point source will be achieved. To align with the overall engine performance calculation method, the turbine coolant point source injection flow rate percentage is given as a percentage of the compressor inlet flow rate (e.g., a coolant flow rate percentage of 1.5% means the coolant flow rate accounts for 1.5% of the total compressor inlet flow rate), and the coolant temperature injected by the coolant point source is given according to the temperature index of different bleed air locations. For example, 1 represents selecting the first temperature value, and 2 represents selecting the second temperature value.

[0068] 4.1 Loading of the airflow direction of the cold air point source on the blade surface.

[0069] Based on the geometric position of the cold air point source on the blade surface set in step 2, the cold air injection direction at that point is set. From the perspective of the global coordinate system, the injection direction of the cold air point source on the blade surface varies along the blade arc length (from the leading edge to the trailing edge). However, the defined input airflow angle is the local coordinate-defined injection angle of the cold air point source at that point (including the circumferential angle α and the radial angle β). The same injection angle at different local coordinate systems may differ in the global coordinate system. Therefore, a built-in coordinate system transformation algorithm can automatically convert the angle parameters in the local coordinate system into the injection direction vector in the global coordinate system, decomposing it into components in the X, Y, and Z directions. This ultimately achieves precise loading of the cold air point source injection direction at any arc length position on the blade surface, and the injection direction at different relative arc length positions can be independently customized to meet diverse design requirements.

[0070] Based on the four defined feature locations, locate the positions defining the air jet direction parameters in the Input.dat file, which include the circumferential angle α and the radial angle β. α is the angle between the local arc length tangent and the air jet direction; α=90° represents the air jet direction being perpendicular to the local blade arc tangent. β is the radial deflection angle of the meridional radial airflow; β=+45° represents a 45° deflection towards the blade tip, and β=-45° represents a 45° deflection towards the blade root.

[0071] 4.2 Loading of cold air jet direction at the upper and lower end areas of the blade wall.

[0072] The loading method for the airflow injection direction of the cold air point source on the upper and lower end areas of the blade is similar to that of the airflow injection direction on the blade surface. The only difference is that the airflow injection direction in the upper and lower end areas of the blade is defined as an angle in the global coordinate system, thus eliminating the need for a transformation from local to global coordinates. Based on the seven characteristic positions defined in the upper and lower end areas of the blade, the location of the airflow injection direction parameter definition in the Input.dat file is found, which includes the circumferential angle α and the radial angle β. At this point, the corresponding angles are the airflow injection angles α = 0 and β = 60° in the global coordinate system. The program incorporates automatic direction vector matching logic based on the difference in airflow angle direction between the upper and lower end areas of the turbine blade: for the lower end area wall cold air point, when the airflow is injected into the inner channel, the radial angle is set to positive, and the program outputs a positive radial component of the airflow injection direction vector; for the upper end area wall cold air point, when the airflow is injected into the inner channel, the radial angle is set to negative, and the program outputs a negative radial component of the airflow injection direction vector. Simultaneously, a cold air point source can be defined at any position along the axial direction of the blade inlet and outlet. The program supports independent setting of the injection direction for different axial positions, ensuring the compatibility of the cold air injection direction in the upper and lower end zones with the airflow characteristics of the flow channel. Through this file, users can define the injection angle corresponding to the axial position. The program will provide intelligent suggested values ​​based on the differences between the upper and lower end zones (the radial component of the lower end zone is positive by default, and that of the upper end zone is negative), thus completing the loading of the cold air point source airflow direction for the end zone blades.

[0073] Step 5: Verify and export the parameters of the air conditioning point source.

[0074] After completing the above four steps, the settings for the "Input.dat" file are finished. This generates two files: "Cooling_Point_Position.dat" and "Cooling_Point_Vector.dat". The "Cooling_Point_Position.dat" file stores the set geometric position (X, Y, Z) coordinates of the turbine blade cooling air. This file is in Tecplot format. The "Cooling_Point_Position.dat" and "Cooling_Point_Vector.dat" files can be imported into the Tecplot program to import the cooling air coordinate point position information. At this point, the Tecplot program can display the turbine blade cooling air point source position and check whether it meets the initial cooling hole setting requirements.

[0075] In addition, the air injection flow rate and air injection angle can be adaptively adjusted based on the engine operating status.

[0076] A pre-defined knowledge base containing multi-dimensional data on state, environment, and cooling is established. Each engine operating state (e.g., takeoff, cruise) is mapped to a defined set of engine operating parameters (e.g., engine speed, turbine inlet temperature). These operating parameters are further uniquely associated with the thermal boundary conditions (blade heat load distribution) and aerodynamic flow environment (mainstream velocity, pressure gradient, secondary flow intensity) of the turbine components under that state. When an aircraft is identified as operating in a specific engine state, the corresponding thermal and flow environment is quickly retrieved through table lookup or model analysis. This forms the input benchmark for dynamic adjustment.

[0077] Using a preset reference state (typically cruising) as a baseline, incremental and proportional adjustments are made based on the environmental differences between the current and reference states. The relative percentage increase or decrease in heat load for each region (e.g., leading edge, blade base) is calculated by comparing the heat flux density distribution on the blade surface between the current and reference states. Subsequently, the flow weights allocated to each cooling air source in that region are adjusted according to this percentage trend, achieving adaptive distribution of more cooling air to areas with high heat loads. The local mainstream velocity vector and end-zone secondary flow intensity are compared between the current and reference states. Based on the physical laws governing the interaction between the film cooling jet and the mainstream, the injection angle is fine-tuned compensatorily. For example, when the mainstream velocity increases, the injection angle is appropriately reduced to bring the jet closer to the wall surface.

[0078] Flow rate adjustment primarily responds to changes in heat load, aiming to ensure that the metal temperature does not exceed limits. Angle adjustment primarily responds to changes in the flow field environment, ensuring effective adhesion and coverage of the cooling air. By comparing the heat flux density distribution on the blade surface in the current state with that in the reference state, the relative percentage increase or decrease in heat load for each region (such as the leading edge and blade flank) is calculated. Subsequently, the flow rate weights allocated to each cooling air source in that region are adjusted according to this percentage trend, achieving adaptive allocation where "regions with high heat loads receive more cooling air." The local mainstream velocity vector and end-zone secondary flow intensity are compared between the current state and the reference state. Based on the physical laws governing the interaction between the film cooling jet and the mainstream (such as jet anti-blow-off capability and film coverage efficiency), the injection angle is finely adjusted for compensation. For example, when the mainstream velocity increases, the injection angle is appropriately reduced to bring the jet closer to the wall surface.

[0079] Adaptive adjustments to the air injection flow rate and angle ensure that the blades are in a safe and near-optimal cooling state under all critical operating conditions. Dynamically enhanced cooling under high heat load conditions guarantees safety; dynamically reduced cooling under low heat load conditions improves overall engine thermal efficiency, ultimately maximizing cooling performance throughout the engine's lifespan.

[0080] In this embodiment, a computer device is provided, such as... Figure 4As shown, it includes a memory 401, a processor 402, and a computer program stored in the memory and executable on the processor. When the processor executes the computer program, it implements the above-mentioned parameterization method for the cooling gas point source of any aero-engine turbine blade.

[0081] Specifically, the computer device can be a computer terminal, a server, or a similar computing device.

[0082] In this embodiment, a computer-readable storage medium is provided, which stores a computer program that executes any of the above-described methods for parameterizing the cooling gas point source of aero-engine turbine blades.

[0083] Specifically, computer-readable storage media include both permanent and non-permanent, removable and non-removable media, which can store information using any method or technology. Information can be computer-readable instructions, data structures, modules of programs, or other data. Examples of computer-readable storage media include, but are not limited to, phase-change memory (PRAM), static random access memory (SRAM), dynamic random access memory (DRAM), other types of random access memory (RAM), read-only memory (ROM), electrically erasable programmable read-only memory (EEPROM), flash memory or other memory technologies, CD-ROM, digital versatile optical disc (DVD) or other optical storage, magnetic tape, magnetic disk storage or other magnetic storage devices, or any other non-transferable medium that can be used to store information accessible by a computing device. As defined herein, computer-readable storage media do not include transient computer-readable media, such as modulated data signals and carrier waves.

[0084] Based on the same inventive concept, this invention also provides a parameterization device for the cold gas point source of an aero-engine turbine blade, as described in the following embodiments. Since the principle behind the problem-solving of the aero-engine turbine blade cold gas point source parameterization device is similar to that of the aero-engine turbine blade cold gas point source parameterization method, the implementation of the aero-engine turbine blade cold gas point source parameterization device can refer to the implementation of the aero-engine turbine blade cold gas point source parameterization method, and repeated details will not be elaborated further. As used below, the terms "unit" or "module" can refer to a combination of software and / or hardware that performs a predetermined function. Although the device described in the following embodiments is preferably implemented in software, hardware implementation, or a combination of software and hardware, is also possible and contemplated.

[0085] Figure 5 This is a structural block diagram of a device for determining the cold gas point source of an aero-engine turbine blade according to an embodiment of the present invention, such as... Figure 5As shown, it includes: turbine blade parameter setting module 501, cold air point source geometry position setting module 502, injection direction setting module 503 and cold air injection parameter setting module 504. The structure is described below.

[0086] The turbine blade parameter setting module 501 is used to initialize the input configuration file, set the attribute parameters of the turbine blade, and save the attribute parameters to the input configuration file.

[0087] The cooling gas point source geometric position setting module 502 is used to perform geometric analysis on the blade shape format of the turbine blade, divide the positioning dimension of the turbine blade into the blade surface feature position and the blade upper and lower end area feature position, and determine the three-dimensional coordinates of the blade cooling gas point source based on the blade geometry file and the positioning dimension.

[0088] The injection direction setting module 503 is used to set the cold air injection angle corresponding to each of the blade cold air point sources, and save the three-dimensional coordinates of the blade cold air point sources and the cold air injection angle to the input configuration file, wherein the cold air injection angle includes a circumferential angle. α With radial angle β ;

[0089] The cold air injection parameter setting module 504 is used to set the cold air injection flow rate and cold air injection temperature of each of the cold air point sources of the blades, and save the cold air injection flow rate and the cold air injection temperature to the input configuration file respectively.

[0090] In one embodiment, the turbine blade parameter setting module includes:

[0091] The turbine blade parameter setting unit is used to set the attribute parameters, including the number of blade rows, blade type, blade data unit, and number of blade channels.

[0092] In one embodiment, the cooling air point source geometry setting module includes:

[0093] The geometric entity recognition unit is used to identify the geometric entities of the turbine blade based on the blade geometry file, wherein the geometric entities include the leading edge line, trailing edge line, suction surface, pressure surface, hub surface, casing surface, and the flange connecting the blade and the end wall.

[0094] The feature position determination unit is used to determine the blade feature position based on the cooling requirements of the turbine blade, wherein the blade feature position includes the blade surface feature position and the blade upper and lower end area feature position.

[0095] A position mapping unit is used to map the blade surface feature positions and the blade upper and lower end area feature positions to the geometric entity corresponding to the blade geometry file;

[0096] The positioning logic definition unit is used to define parameterized positioning logic for the mapped blade surface feature positions and the blade upper and lower end area feature positions, respectively. The positioning logic for the blade surface feature positions includes dimensionless arc length and spanwise distribution parameters, and the positioning logic for the blade upper and lower end area feature positions includes end area anchor point coordinates and phase angle range.

[0097] The blade surface three-dimensional coordinate calculation unit is used to calculate the three-dimensional coordinates of multiple cold air point sources on the blade surface using the dimensionless arc length and the spanwise distribution parameters.

[0098] The blade end region three-dimensional coordinate calculation unit is used to calculate the three-dimensional coordinates of multiple cold air point sources in the upper and lower end regions of the blade using the end region anchor point coordinates and the phase angle range.

[0099] In one embodiment, the feature location determination unit is further configured to determine a typical region of thermal load on the turbine blade surface, wherein the typical region of thermal load includes a leading edge region, a pressure-side blade head region, and a trailing edge region; based on the typical region of thermal load, the surface of the turbine blade is divided into multiple blade surface feature locations, wherein the blade surface feature locations include a suction-side leading edge, a pressure-side leading edge, a pressure-side blade head, and a trailing edge; the end region cooling area and the tip clearance heat transfer area requiring cooling in the connection end region between the blade and the hub are determined, and based on the end region cooling area and the tip clearance heat transfer area, the upper and lower end regions of the blade are divided into multiple upper and lower end region feature locations, wherein the upper and lower end region feature locations include a hub blade inter-blade edge plate, a casing blade inter-blade edge plate, a hub blade leading edge plate, a casing blade leading edge plate, a hub blade trailing edge plate, a casing blade trailing edge plate, and a moving blade tip region.

[0100] In one embodiment, the injection direction setting module includes:

[0101] The surface injection angle setting unit is used to set the radial angle if the blade cooling air point source is a blade surface cooling air point source. β and circumferential angle α As the angle of the cold air injection;

[0102] The end-area injection angle setting unit is used to adjust the radial angle of the cold air injection angle based on the end-area positions of the upper and lower end-area cold air source of the blade, if the blade's cold air source is located in the upper or lower end-area cold air source. β Make corrections, and adjust the radial angle accordingly. β and the set circumferential angle α As the cold air injection angle, the end region position includes the upper end of the blade and the lower end of the blade.

[0103] In one embodiment, the end zone injection angle setting unit is further configured to adjust the radial angle if the cool air point source in the upper and lower end zones of the blade is located at the lower end of the blade. β Set to positive; if the cold air source in the upper and lower end regions of the blade is located at the upper end of the blade, set the radial angle to positive. β Set to negative.

[0104] In one embodiment, the above-described apparatus further includes an adaptive adjustment module.

[0105] In one embodiment, the adaptive adjustment module includes:

[0106] The parameter setting unit is used to define multiple engine operating states and a set of engine operating condition parameters corresponding to each engine operating state in the input configuration file, wherein the engine operating states include takeoff state, cruise state and maneuver state;

[0107] The parameter set and flow field information acquisition unit is used to determine the blade thermal boundary condition parameter set and mainstream flow field information under the current engine operating state from a preset association database based on the current engine operating state and the engine operating condition parameter set corresponding to the current engine operating state. The blade thermal boundary condition parameter set includes the heat load reference distribution of the characteristic positions on the blade surface and the characteristic positions at the upper and lower end regions of the blade.

[0108] The cold air injection flow rate adjustment unit is used to compare the blade thermal boundary condition parameter set of the current engine operating state with the blade thermal boundary condition parameter set of the preset reference state, calculate the relative change of cooling demand at the characteristic positions of the blade surface and the characteristic positions of the upper and lower end regions of the blade, and adjust the cold air injection flow rate of each blade cold air point source in the input configuration file according to the relative change.

[0109] The cold air injection angle correction unit is used to adjust the circumferential angle of the cold air injection angle of the cold air point source on the blade surface based on the mainstream flow field information under the current engine operating condition. α With radial angle β Make corrections;

[0110] The configuration file update unit is used to update the adjusted air injection flow rate and the corrected air injection angle into the input configuration file.

[0111] Compared with existing turbine blade cooling gas point source loading methods, the cooling gas point source parameterization method of this invention achieves significant improvements in efficiency, time consumption, and accuracy:

[0112] 1. Improved efficiency: The loading speed is increased by two orders of magnitude compared to traditional methods, significantly shortening the design cycle of turbine blade cooling efficiency.

[0113] 2. Reduced time consumption: The time required for manual operation has been reduced from the traditional 5 hours to within 5 seconds, greatly reducing the time cost for designers.

[0114] 3. Precision optimization: Through parameterized automatic loading and algorithm verification, human operation errors are avoided, the error rate is reduced by 99%, and the accuracy of air conditioning point source loading and design reliability are guaranteed.

[0115] 4. Strong adaptability and compatibility: Supports mainstream formats such as COORD and Geomturbo, adapts to stator / rotor and high-pressure / low-pressure turbine blades, and can cover 4 types of feature positions on the blade surface and 7 types of end regions without the need to redevelop logic for different blades.

[0116] 5. Iteration convenience: When the blade geometry or cooling parameters are modified, only the corresponding parameters in the "Input.dat" file need to be updated, and the program will automatically recalculate and generate the results without having to repeat the entire process.

[0117] The embodiments of the present invention achieve the following technical effects:

[0118] This invention provides a parameterization method for cold gas point sources on aero-engine turbine blades. It is applicable to the precise positioning, orientation setting, and automatic parameter loading of cold gas point sources on the surface and upper and lower end walls of turbine blades. It can directly provide high-precision cold gas boundary conditions for turbine blade cooling effect simulation, cooling scheme optimization, and turbine performance evaluation, and is particularly suitable for cooling design scenarios of multi-bladed turbine blades with complex flow channels. The method constructs a complete parameterization system encompassing "model analysis - parameter definition - intelligent conversion - verification and export," achieving automatic analysis compatible with mainstream blade model formats without manual geometric identification. It includes parameterized definition and automatic calculation of cold gas point source position, direction, temperature, and flow rate; automatic conversion between local and global coordinate systems to avoid non-physical errors in injection direction; and batch loading and rapid iteration of multi-bladed and multi-feature positions, significantly shortening the design cycle. Visual verification ensures parameter accuracy, reducing human error.

[0119] Obviously, those skilled in the art should understand that the modules or steps of the above-described embodiments of the present invention can be implemented using general-purpose computing devices. They can be centralized on a single computing device or distributed across a network of multiple computing devices. Optionally, they can be implemented using computer-executable program code, thereby storing them in a storage device for execution by a computing device. In some cases, the steps shown or described can be performed in a different order than those presented here, or they can be fabricated as separate integrated circuit modules, or multiple modules or steps can be fabricated as a single integrated circuit module. Thus, the embodiments of the present invention are not limited to any particular hardware and software combination.

[0120] The above description is merely a preferred embodiment of the present invention and is not intended to limit the present invention. For those skilled in the art, various modifications and variations can be made to the embodiments of 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. A method for determining the point source of cooling gas for aero-engine turbine blades, characterized in that, include: Initialize the input configuration file, set the attribute parameters of the turbine blades, and save the attribute parameters to the input configuration file; Geometric analysis is performed on the blade profile of the turbine blade, and the positioning dimension of the turbine blade is divided into the blade surface feature position and the blade upper and lower end area feature position. Based on the blade geometry file, the three-dimensional coordinates of the blade cooling gas point source are determined according to the positioning dimension. The blade cooling gas point source includes the blade surface cooling gas point source and the blade upper and lower end area cooling gas point source. The positioning dimensions of the turbine blade are divided into blade surface feature positions and blade upper and lower end region feature positions. Based on the blade geometry file, the three-dimensional coordinates of the blade cooling gas point source are determined according to the positioning dimensions, including: Based on the blade geometry file, the geometric entities of the turbine blade are identified, wherein the geometric entities include the leading edge line, trailing edge line, suction surface, pressure surface, hub surface, casing surface, and the flange connecting the blade and the end wall. Based on the cooling requirements of turbine blades, the characteristic locations of the blades are determined, including the characteristic locations on the blade surface and the characteristic locations at the upper and lower ends of the blade. Map the blade surface feature positions and the blade upper and lower end region feature positions to the geometric entities corresponding to the blade geometry file; Parametric positioning logic is defined for the mapped blade surface feature positions and the blade upper and lower end region feature positions, respectively. The positioning logic for the blade surface feature positions includes dimensionless arc length and spanwise distribution parameters, and the positioning logic for the blade upper and lower end region feature positions includes end region anchor point coordinates and phase angle range. The three-dimensional coordinates of multiple cold air point sources on the blade surface are calculated using the dimensionless arc length and the spanwise distribution parameters. The three-dimensional coordinates of multiple cold air point sources in the upper and lower end regions of the blades are calculated using the end region anchor point coordinates and the phase angle range. Set the air injection angle corresponding to each of the blade air source points, and save the three-dimensional coordinates of the blade air source points and the air injection angle to the input configuration file. The air injection angle includes the circumferential angle. α With radial angle β ; Set the air injection flow rate and air injection temperature for each of the blade air source points, and save the air injection flow rate and air injection temperature to the input configuration file respectively.

2. The method for determining the cold gas point source of an aero-engine turbine blade as described in claim 1, characterized in that, Based on the cooling requirements of the turbine blades, the characteristic locations of the blades are determined, including: The typical regions of thermal load on the turbine blade surface are determined, wherein the typical regions of thermal load include the leading edge region, the pressure-side blade basin region, and the trailing edge region. Based on the typical region of the heat load, the surface of the turbine blade is divided into multiple blade surface feature locations, wherein the blade surface feature locations include the suction-side leading edge, the pressure-side leading edge, the pressure-side blade head, and the trailing edge. The cooling areas of the end region that need cooling and the heat transfer area of ​​the blade tip gap that need cooling are determined in the connection area between the blade and the hub. Based on the cooling areas of the end region and the heat transfer area of ​​the blade tip gap, the upper and lower end regions of the blade are divided into multiple characteristic positions of the upper and lower end regions of the blade. The characteristic positions of the upper and lower end regions of the blade include the hub blade inter-edge plate, the casing blade inter-edge plate, the hub blade leading edge plate, the casing blade leading edge plate, the hub blade trailing edge plate, the casing blade trailing edge plate, and the moving blade tip region.

3. The method for determining the cold gas point source of an aero-engine turbine blade as described in claim 1, characterized in that, Setting the cold air injection angle corresponding to each of the blade cold air point sources includes: If the blade cooling air source is a cooling air source on the blade surface, the set radial angle β and circumferential angle α As the angle of the cold air injection; If the cold air source for the blade is located in the upper or lower end region of the blade, the radial angle of the cold air injection angle is determined based on the end region location of the cold air source in the upper or lower end region of the blade. β Make corrections, and adjust the radial angle accordingly. β and the set circumferential angle α As the cold air injection angle, the end region position includes the upper end of the blade and the lower end of the blade.

4. The method for determining the cold gas point source of an aero-engine turbine blade as described in claim 3, characterized in that, Based on the location of the cold air point source at the upper and lower ends of the blade, the radial angle of the cold air injection angle is determined. β Make corrections, including: If the cold air source in the upper and lower end regions of the blade is located at the lower end of the blade, the radial angle will be... β Set to positive; If the cold air source in the upper and lower end regions of the blade is located at the upper end of the blade, the radial angle will be... β Set to negative.

5. The method for determining the cold gas point source of an aero-engine turbine blade as described in claim 1, characterized in that, The attribute parameters include: Number of blade rows, blade type, blade data unit, and number of blade channels.

6. The method for determining the cold gas point source of an aero-engine turbine blade as described in any one of claims 1 to 5, characterized in that, Also includes: Adaptive adjustment of the cold air injection flow rate and the cold air injection angle based on the engine operating status includes: The input configuration file defines multiple engine operating states and a set of engine operating condition parameters corresponding to each engine operating state, wherein the engine operating states include takeoff state, cruise state and maneuver state; Based on the current engine operating state and the engine operating condition parameter set corresponding to the current engine operating state, the blade thermal boundary condition parameter set and mainstream flow field information under the current engine operating state are determined from a preset association database. The blade thermal boundary condition parameter set includes the heat load reference distribution of the characteristic positions on the blade surface and the characteristic positions at the upper and lower end regions of the blade. By comparing the set of blade thermal boundary condition parameters under the current engine operating state with the set of blade thermal boundary condition parameters under the preset reference state, the relative changes in cooling requirements of the blade surface feature positions and the blade upper and lower end region feature positions are calculated, and the cold air injection flow rate of each blade cold air point source in the input configuration file is adjusted according to the relative changes. Based on the mainstream flow field information under the current engine operating conditions, the circumferential angle of the cold air injection angle of the cold air point source on the blade surface is determined. α With radial angle β Make corrections; Update the adjusted air jet flow rate and corrected air jet angle to the input configuration file.

7. A device for determining the point source of cooling gas for aero-engine turbine blades, characterized in that, include: The turbine blade parameter setting module is used to initialize the input configuration file, set the attribute parameters of the turbine blade, and save the attribute parameters to the input configuration file. The cooling gas point source geometric position setting module is used to perform geometric analysis on the blade shape format of the turbine blade, divide the positioning dimension of the turbine blade into the blade surface feature position and the blade upper and lower end area feature position, and determine the three-dimensional coordinates of the blade cooling gas point source based on the blade geometry file and the positioning dimension. The air conditioning point source geometry setting module includes: The geometric entity recognition unit is used to identify the geometric entities of the turbine blade based on the blade geometry file, wherein the geometric entities include the leading edge line, trailing edge line, suction surface, pressure surface, hub surface, casing surface, and the flange connecting the blade and the end wall. The feature position determination unit is used to determine the blade feature position based on the cooling requirements of the turbine blade, wherein the blade feature position includes the blade surface feature position and the blade upper and lower end area feature position. A position mapping unit is used to map the blade surface feature positions and the blade upper and lower end area feature positions to the geometric entity corresponding to the blade geometry file; The positioning logic definition unit is used to define parameterized positioning logic for the mapped blade surface feature positions and the blade upper and lower end area feature positions, respectively. The positioning logic for the blade surface feature positions includes dimensionless arc length and spanwise distribution parameters, and the positioning logic for the blade upper and lower end area feature positions includes end area anchor point coordinates and phase angle range. The blade surface three-dimensional coordinate calculation unit is used to calculate the three-dimensional coordinates of multiple cold air point sources on the blade surface using the dimensionless arc length and the spanwise distribution parameters. The blade end region three-dimensional coordinate calculation unit is used to calculate the three-dimensional coordinates of multiple cold air point sources in the upper and lower end regions of the blade using the end region anchor point coordinates and the phase angle range. The injection direction setting module is used to set the cold air injection angle corresponding to each of the blade cold air point sources, and save the three-dimensional coordinates of the blade cold air point sources and the cold air injection angle to the input configuration file, wherein the cold air injection angle includes the circumferential angle. α With radial angle β ; The cold air injection parameter setting module is used to set the cold air injection flow rate and cold air injection temperature for each of the cold air point sources on the blades, and save the cold air injection flow rate and the cold air injection temperature to the input configuration file respectively.

8. A computer device, comprising a memory, a processor, and a computer program stored in the memory and executable on the processor, characterized in that, When the processor executes the computer program, it implements the method for determining the cold air point source of the aero-engine turbine blade as described in any one of claims 1 to 6.

9. A computer-readable storage medium, characterized in that, The computer-readable storage medium stores a computer program that performs the method for determining the cold gas point source of an aero-engine turbine blade according to any one of claims 1 to 6.

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