A turbine guide vane and its throat area design and adjustment method

Abstract

The application provides a turbine guide vane and a throat area design and adjustment method thereof, and the method comprises the following steps: turbine guide vane aerodynamic and cooling structure design is performed; a three-dimensional geometric model of at least one cascade passage is established, the pressure and temperature of inlet gas and inlet cold gas are given, gas flow calculation is performed, the gas flow and a Mach surface in a hot cascade are obtained; the theoretical hot state throat area of the turbine guide vane is calculated; the heat insulation coating design of the turbine guide vane blade body and the rim plate flow channel surface is performed; the coating process test of the turbine guide vane is performed, the actual coating thickness of each characteristic position point is measured, the actual coating thickness is used to calculate the throat area of the turbine guide vane after the heat insulation coating is sprayed; the turbine guide vane is rotated at different angles around the blade radial direction, and the gas flow of the cascade passage under different angles is obtained; when the flow and the throat area meet the requirements, the throat area of the turbine guide vane with the coating under the rotation angle meets the gas flow requirement.

Description

Technical Field

[0001] This application belongs to the field of aero-engine technology, and specifically relates to a turbine guide vane and its throat area design and adjustment method. Background Technology

[0002] For turbine blades in aero-engines, the narrowest part of the guide vane cascade has the smallest flow area and the greatest restriction on airflow. Therefore, the size of the guide vane throat area directly affects the overall engine performance. Due to various reasons such as the application of thermal barrier coatings, assembly clearances, and measurement methods, it is difficult to achieve a guide vane throat area within the design requirements during actual machining. Therefore, it is necessary to design a method for designing and adjusting the blade throat area with thermal barrier coatings.

[0003] Current turbine guideway throat area measurements are based on the cold-state throat area, not the hot-state aerodynamic throat area. The cold and hot throat areas differ in both value and location, resulting in a discrepancy between the actual manufactured guideway area and the required area.

[0004] In addition, current guide vanes often use different groups of blades to adjust in order to ensure the final guide vane area. However, this requires the processing of multiple sets of molds and their blades, which greatly increases production costs. Furthermore, different groups of blades have different outlet areas, resulting in different gas flow rates. This can cause complex unsteady excitation effects on the rotor blades downstream, potentially causing significant fluctuations in the wake excitation force within a short period of time. Summary of the Invention

[0005] The purpose of this application is to provide a turbine guide and a method for designing and adjusting its throat area, so as to solve or alleviate at least one of the problems in the prior art.

[0006] The technical solution of this application is: a method for designing and adjusting the throat area of ​​a turbine guide vane, the method comprising:

[0007] The aerodynamic and cooling structure design of the turbine guide vanes was carried out to obtain the blade design parameters;

[0008] Based on the blade design parameters, establish a three-dimensional geometric model of at least one blade cascade channel. In the three-dimensional geometric model, given the inlet gas pressure and temperature, and the cold air inlet pressure and temperature, calculate the gas flow rate M0 and the one-time speed surface in the hot blade cascade. The intersection line between the one-time speed surface and the back side of the blade cascade channel is called the characteristic line.

[0009] Calculate the theoretical hot throat area F0 of the turbine guide vane;

[0010] Design the thermal insulation coating for the blade body and rim flow channel of the turbine guide vane, and determine the parameters of the thermal insulation coating;

[0011] Coating process experiments were conducted on turbine guide vanes under thermal insulation coating parameters, and the actual coating thickness at various characteristic locations was measured under these parameters.

[0012] Calculate the throat area F1 of the turbine guide vane after the thermal insulation coating is applied based on the actual coating thickness.

[0013] By rotating the turbine guide vanes radially around the blades at different angles, the gas flow rate M1 in the cascade passage at different angles was obtained;

[0014] When (M1-M0) / M0≈(F1-F0) / F0, the throat area of ​​the coated turbine guide vane at this rotation angle meets the gas flow requirements.

[0015] In a preferred embodiment of this application, the process of calculating the theoretical hot throat area F0 of the turbine guide vane is as follows:

[0016] Establish a cylindrical coordinate system with the engine axis X as the axis. Intersect the characteristic line with n sets of cylindrical tangents of different radii. Calculate the average value of the distance line W between the intersection points on the back side of each set. Establish a straight line by connecting the midpoints of the line segments with the largest and smallest radii. The distance between the two intersection points of the straight line and the upper and lower edge flow channels is the height line H. Then, the theoretical hot throat area of ​​the turbine guide vane is F0 = W * H. The intersection points of the height line H and the distance line W are collectively referred to as characteristic position points.

[0017] In a preferred embodiment of this application, the number of cylinder groups n is 3 to 7.

[0018] In a preferred embodiment of this application, the method for calculating the throat area F1 of the turbine guide vane after applying the heat-insulating coating is as follows:

[0019] F1 = F0 - H(a+b) - W(c+d)

[0020] In the formula, a and b are the average coating thicknesses at characteristic locations on the basin and back sides of the turbine guide vane, respectively.

[0021] c and d are the average coating thicknesses at characteristic locations on the flow channel surfaces of the upper and lower edge plates, respectively.

[0022] In a preferred embodiment of this application, the method further includes:

[0023] If the actual throat area of ​​the turbine guide vane does not meet the actual flow rate requirement of the whole machine after production and processing, the coating thickness is adjusted so that the throat area of ​​the turbine guide vane meets the required flow rate M2.

[0024] In a preferred embodiment of this application, the process of adjusting the coating thickness to ensure that the throat area of ​​the turbine guide vane meets the required flow rate M2 is as follows:

[0025] Adjust the coating thickness of the turbine guide vane on the basin side, back side, upper edge plate flow channel surface, and lower edge plate flow channel surface. Record the adjustment values ​​as △a, △b, △c, and △d, respectively. Calculate the adjusted turbine guide vane throat area F2.

[0026] When (M2-M1) / M1≈(F2-F1) / F1, the coating thickness at this point meets the actual flow rate required by the whole machine.

[0027] In a preferred embodiment of this application, the adjusted turbine guide throat area F2 is calculated as follows:

[0028] F2=F1-H(△a+△b)-W(△c+△d).

[0029] In addition, this application also provides a turbine guide, which is designed according to any of the turbine guide throat area design and adjustment methods described above.

[0030] The turbine guide vane throat area design and adjustment method proposed in this application takes into account the effects of thermal aerodynamics and thermal barrier coating on the throat, which can ensure that the guide vane area meets the gas flow requirements, and ensure the consistency of the blade shape and channel outlet area of ​​each guide vane in the guide vane, while saving costs. Attached Figure Description

[0031] To more clearly illustrate the technical solutions provided in this application, the accompanying drawings will be briefly described below. Obviously, the drawings described below are merely some embodiments of this application.

[0032] Figure 1 A flowchart illustrating the method for designing and adjusting the back passage area of ​​the turbine guide vane in this application.

[0033] Figure 2 This is a cross-sectional Ma distribution diagram of a turbine guide in an embodiment of this application.

[0034] Figure 3 This is a schematic diagram showing the intersection of three sets of cylindrical tangents of different radii with the blade according to an embodiment of this application. Detailed Implementation

[0035] To make the objectives, technical solutions, and advantages of this application clearer, the technical solutions in the embodiments of this application will be described in more detail below with reference to the accompanying drawings.

[0036] This application proposes a method for designing and adjusting the throat area of ​​a turbine guide vane. By considering the effects of thermal aerodynamics and thermal barrier coating on the throat, the guide vane area can be guaranteed to meet the gas flow requirements, ensuring the consistency of the blade shape and channel outlet area of ​​each guide vane in the guide vane.

[0037] like Figure 1As shown, the turbine guideway throat area design and adjustment method provided in the application includes the following process:

[0038] S1. Perform aerodynamic and cooling structure design of turbine guide vanes to obtain blade design parameters.

[0039] S2. Establish a three-dimensional geometric model of the blade cascade passage based on the blade design parameters. In the three-dimensional geometric model, given the inlet gas pressure and temperature, and the inlet pressure (or cold air flow rate) and temperature, perform gas flow calculations to obtain the gas flow rate M0 and the 1-times-sonic surface within the hot blade cascade, such as... Figure 2 As shown, the intersection of the one-time speed surface and the back side of the blade channel is called the characteristic line, and the arrow is the cutoff line of the speed surface.

[0040] S3. Calculate the theoretical hot throat area F0 of the turbine guide vane. The calculation method is as follows:

[0041] like Figure 3 As shown, a cylindrical coordinate system is established with the direction of the engine axis X as the axis. The characteristic line is intersected by n sets of cylindrical tangents with different radii (n takes 3 to 7). The distance line W between the intersection points on the back side of each set is calculated. The average value of the distance line W is taken. A straight line is established by connecting the midpoint of the line segment with the largest and smallest radii. The two intersection points of the straight line with the flow channel surface of the edge plate form the height line H. Then F0 = W * H. The intersection points of the height line H and the distance line W are collectively referred to as characteristic position points.

[0042] S4. Design the heat insulation coating for the turbine guide vane blade body and rim flow channel surface, and determine the heat insulation coating parameters.

[0043] S5. Conduct coating process tests on turbine guide vanes under the thermal insulation coating parameters, and measure the actual coating thickness at each characteristic location point under the thermal insulation coating parameters. The average coating thickness at the characteristic location points on the basin side and back side of the turbine guide vane is recorded as a and b, respectively, and the average coating thickness at the characteristic location points on the upper edge plate and lower edge plate is recorded as c and d, respectively.

[0044] The calculation method for the throat area F1 of the turbine guide vane after the thermal insulation coating is applied is as follows:

[0045] F1 = F0 - H(a+b) - W(c+d).

[0046] S6. Using the theoretical model, rotate it around the engine z-axis (blade radial direction) at different angles, and repeat step 2 to obtain the gas flow rate M1 of the blade passage at different angles.

[0047] S7. When (M1-M0) / M0≈(F1-F0) / F0, record the rotation angle at this time. At this rotation angle, the throat area of ​​the coated turbine guide vane meets the gas flow requirements and can be used for production and processing.

[0048] Table 1 shows the design parameters for the throat area of ​​a high-pressure turbine guide vane with a thermal barrier coating, obtained using steps S1 to S7.

[0049] Table 1 Design Parameters

[0050]

[0051]

[0052] In addition, the method of this application also includes:

[0053] S8. After production and processing, if the actual machined throat area of ​​the turbine guide vanes does not meet the actual flow rate requirement of the entire machine due to changes in overall machine requirements or processing reasons, it can be adjusted by adjusting the coating thickness according to the required flow rate M2, that is:

[0054] Adjust the coating thickness of the turbine guide vane on the basin side, back side, upper edge plate flow channel surface, and lower edge plate flow channel surface. Record the adjustment values ​​as △a, △b, △c, and △d, respectively. Calculate the throat area F2 after adjustment. The calculation formula is F2=F1-H(△a+△b)-W(△c+△d).

[0055] When (M2-M1) / M1≈(F2-F1) / F1, the coating thickness at this point meets the actual requirements of the whole machine.

[0056] For example, in the following embodiment of this application, the throat area of ​​a low-pressure turbine guide vane is too small, resulting in a 4% reduction in gas flow rate. The coating thickness on the blade basin side, back side, upper edge plate flow channel surface, and lower edge plate flow channel surface is increased according to step S8 of the above method. Testing shows that the turbine guide vane flow rate meets the requirements.

[0057] Table 2 Effects of adjusting the throat area of ​​blades and guide vanes

[0058]

[0059] The turbine guide vane throat area design and adjustment method proposed in this application takes into account the effects of thermal aerodynamics and thermal barrier coating on the throat, which can ensure that the guide vane area meets the gas flow requirements, and ensure the consistency of the blade shape and channel outlet area of ​​each guide vane in the guide vane, while saving costs.

[0060] The above description is merely a specific embodiment of this application, but the scope of protection of this application is not limited thereto. Any variations or substitutions that can be easily conceived by those skilled in the art within the technical scope disclosed in this application should be included within the scope of protection of this application. Therefore, the scope of protection of this application should be determined by the scope of the claims.

Claims

1. A method for designing and adjusting the throat area of ​​a turbine guide vane, characterized in that, The method includes: The aerodynamic and cooling structure design of the turbine guide vanes was carried out to obtain the blade design parameters; Based on the blade design parameters, establish a three-dimensional geometric model of at least one blade cascade channel. In the three-dimensional geometric model, given the inlet gas pressure and temperature, and the cold air inlet pressure and temperature, calculate the gas flow rate M0 and the one-time speed surface in the hot blade cascade. The intersection line between the one-time speed surface and the back side of the blade cascade channel is called the characteristic line. The theoretical hot throat area F0 of the turbine guide vane is calculated as follows: A cylindrical coordinate system is established with the engine axis X as the axis. The characteristic line is intersected by n sets of cylindrical tangents with different radii. The average value of the distance line W between the intersection points on the back side of each set is calculated. A straight line is established by connecting the midpoint of the line segment with the largest and smallest radii. The distance between the two intersection points of the straight line and the upper and lower edge flow channels is the height line H. Then the theoretical hot throat area F0 of the turbine guide vane is F0 = W * H. The intersection points of the height line H and the distance line W are collectively referred to as characteristic position points. Design the thermal insulation coating for the blade body and rim flow channel of the turbine guide vane, and determine the parameters of the thermal insulation coating; Coating process experiments were conducted on turbine guide vanes under thermal insulation coating parameters, and the actual coating thickness at various characteristic locations was measured under these parameters. The throat area F1 of the turbine guide vane after the heat insulation coating is sprayed is calculated based on the actual coating thickness. F1 = F0 - H(a+b) - W(c+d), where a and b are the average coating thicknesses at characteristic points on the basin side and back side of the turbine guide vane, respectively; c and d are the average coating thicknesses at characteristic points on the flow channel surfaces of the upper and lower edge plates, respectively. By rotating the turbine guide vanes radially around the blades at different angles, the gas flow rate M1 in the cascade passage at different angles was obtained; When (M1-M0) / M0≈(F1-F0) / F0, the throat area of ​​the coated turbine guide vane at this rotation angle meets the gas flow requirements.

2. The method for designing and adjusting the throat area of ​​a turbine guide vane as described in claim 1, characterized in that, The number of cylinder groups n can be 3 to 7.

3. The method for designing and adjusting the throat area of ​​a turbine guide vane as described in claim 1, characterized in that, The method further includes: If the actual throat area of ​​the turbine guide vane does not meet the actual flow rate requirement of the whole machine after production and processing, the coating thickness is adjusted so that the throat area of ​​the turbine guide vane meets the required flow rate M2.

4. The method for designing and adjusting the throat area of ​​a turbine guide vane as described in claim 3, characterized in that, The process of adjusting the coating thickness to ensure the turbine guideway throat area meets the required flow rate M2 is as follows: Adjust the coating thickness of the turbine guide vane on the basin side, back side, upper edge plate flow channel surface, and lower edge plate flow channel surface. Record the adjustment values ​​as △a, △b, △c, and △d, respectively. Calculate the adjusted turbine guide vane throat area F2. When (M2-M1) / M1≈(F2-F1) / F1, the coating thickness at this point meets the actual flow rate required by the whole machine.

5. The method for designing and adjusting the throat area of ​​a turbine guide vane as described in claim 4, characterized in that, The adjusted calculation method for the turbine guideway throat area F2 is as follows: F2=F1-H(△a+△b)-W(△c+△d).

6. A turbine guide, characterized in that, The turbine guide is designed according to the turbine guide throat area design and adjustment method as described in any one of claims 1 to 5.

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