A low consistency gas cooled turbine guide vane airfoil design method and system

Abstract

The application discloses a low-consistency gas-cooled turbine guider blade profile design method and system, and belongs to the technical field of turbine guider blade profile design, wherein the method comprises initial blade profile design, simulation calculation, load partition, blade profile design coupled with gas film holes and blade profile updating. By considering the design of the gas film holes in the blade profile design, the influence of the blade profile and the outflow of the gas film holes at typical positions can be considered under the condition of large cold gas volume, so that the outflow mixing loss of the suction surface gas film holes is greatly reduced, and the aerodynamic performance of the turbine guider is improved.

Description

Technical Field

[0001] This application belongs to the field of turbine guide vane design technology, and specifically relates to a method and system for designing air-cooled turbine guide vanes with low viscosity. Background Technology

[0002] The inlet temperature of gas turbines in high-performance aero engines can reach 1900~2100K, far exceeding the temperature resistance limit of existing blade materials. Therefore, a large volume of cooling air is necessary to cool the turbine blades. The cooling air consumption of the first-stage guide vane blades (hereinafter referred to as the first-stage guide vane) is close to half of the total cooling air volume of the entire engine. In high-cycle-parameter engines, the cooling air flow rate of the first-stage guide vane can reach 10%~15% of the compressor inlet flow rate. To reduce the number of aero engine parts and lower the overall weight, aerodynamic designs are generally adopted, such as the high-load, low-density airfoil disclosed in patent CN113836633A. This design can reduce the number of blades by 1 / 3 compared to conventional density airfoils while increasing the airfoil load. However, this new airfoil completely changes the internal flow field distribution of the blades, making the traditional separate design mode no longer applicable. Specifically, in existing airfoil designs, the airfoil design and film cooling hole design of the guide vane are separate, referring to... Figure 1 a. Under normal operating conditions, the combustion engine adopts a conventional consistency airfoil design. In this case, due to the large number of blades and the fact that most of the suction surface film air holes are located in the low Mach number leading edge region of the throat, coupled with the early low turbine inlet temperature and small cooling gas consumption, the film air outflow has a negligible impact on aerodynamic performance, resulting in low design coupling requirements. However, referring to... Figure 1 b. High-load, low-consistency airfoils alter the blade flow field distribution. Low-consistency fuel-guided blades have a forward-moving throat and the suction surface film air holes are located in a high Mach number flow field, significantly aggravating the mixing loss between the outflowing cold air and the mainstream fuel gas. If the traditional airfoil design approach is still used, the aerodynamic performance is optimized by increasing the load at the front of the blade, further amplifying the cold air mixing loss. At the same time, existing film air hole designs only focus on blade cooling efficiency. To ensure blade cooling and smooth cold air outflow, film air holes are often placed in the throat and downstream high Mach number regions, and are matched with high cold air flow and high exhaust ratio conditions, further aggravating the mixing loss between the mainstream and cold air. In addition, current multi-objective optimization of airfoils with film air holes mostly relies on full three-dimensional optimization methods to carry out airfoil-only optimization or airfoil-film air hole co-optimization. This approach is difficult to mesh, consumes high computing power, has a long development cycle, and cannot meet the rapid simulation optimization needs of the design phase, highlighting the industry's technical pain points. Summary of the Invention

[0003] This invention addresses the problem of significant mixing losses between the cold gas outflow and the mainstream flow on the suction surface of the first-stage guide vanes of large-volume ultra-low-consistency gas turbines (hereinafter referred to as "first-stage guide vanes"). It proposes a high-load, ultra-low-consistency, strong cold gas mixing, and low-loss airfoil design method and system based on airfoil load zoning control. The method considers the influence of cold gas outflow from film cooling holes at typical locations. By implementing zoning control of the airfoil load, the mixing losses between the cooling outflow from the film cooling holes on the suction surface (especially near and behind the throat) and the mainstream flow are significantly reduced without affecting blade cooling. This reduces the total pressure loss of the first-stage guide vane and improves the aerodynamic performance of the gas turbine.

[0004] A method for designing the airfoil profile of a low-consistency air-cooled turbine guide vane, wherein the guide vane has multiple exhaust film holes on its suction surface, the method comprising: Preliminary design of the blade profile for the low-consistency guide vane was carried out to obtain the initial blade profile; The load distribution of the initial airfoil is obtained through simulation calculation, so as to obtain the Mach number distribution curve of the suction surface of the initial airfoil; The blade cascade channel formed by adjacent initial blade profiles is divided into partitions, and the suction surface load distribution curve is divided into multiple regions along the flow direction based on the partition boundaries of the blade cascade channel. The region where the film cooling orifice is located on the Mach number distribution curve of the suction surface is determined. When at least one film cooling orifice is located in the set region of the Mach number distribution curve of the suction surface, the strong cold air mixing effect of the cold air outflow from the film cooling orifice and its influence on the airfoil loss are coupled and considered. The airfoil control parameters in the initial airfoil region where the film cooling orifice is located are adjusted individually or jointly to reduce the Mach number at the location of the film cooling orifice and keep the Mach number distribution curve at other locations unchanged, so that the influence of the cold air outflow on the mixing loss of the mainstream is minimized, and a new airfoil considering the influence of the cold air outflow from the film cooling orifice is obtained. Calculate the total pressure loss coefficient of the updated airfoil. When the total pressure loss coefficient meets the design requirements, the new airfoil is output as the design airfoil.

[0005] Furthermore, the cascade channels formed by adjacent initial airfoils are divided into zones, and the suction surface Mach number distribution curve is divided into multiple regions along the flow direction based on the zone boundaries of the cascade channels, including: Draw a perpendicular line from the initial blade trailing edge tangent point to the suction surface of the adjacent blade to determine the throat position of the blade passageway; The throat position is offset according to the set width to divide the throat area in the blade channel; The upstream portion of the blade passageway from the initial airfoil leading edge to the throat region is divided into the pre-throat region; The downstream portion of the blade passageway from the initial blade shape trailing edge to the throat region is divided into the post-throat region. Determine the boundary points of the pre-throat region, throat region, and post-throat region in the suction surface Mach number distribution curve, and correspondingly divide the suction surface Mach number distribution curve along the flow direction into the pre-throat region, throat region, and post-throat region.

[0006] Furthermore, considering the strong cold air mixing effect of the cold air outflow from the film cooling orifice and its impact on airfoil losses, the airfoil control parameters within the initial airfoil region where the film cooling orifice is located are individually or jointly adjusted to reduce the Mach number at the location of the film cooling orifice while keeping the Mach number distribution curves at other locations unchanged. This minimizes the impact of the cold air outflow on the mixing losses of the mainstream. This includes individually or jointly adjusting the initial airfoil's installation angle, trailing edge bend angle, inlet configuration angle, leading edge wedge angle, outlet configuration angle, and surface curvature. This results in a Mach number at the film cooling orifice location that is 5% to 20% or more lower than the initial airfoil, and a total pressure loss that is 3% or more lower.

[0007] Furthermore, when adjusting the airfoil control parameters, the throat area of ​​the guide vane remains constant.

[0008] Furthermore, it can significantly reduce the mixing loss of the mainstream by the air film pores in the throat region and / or the post-throat region.

[0009] Furthermore, the method also includes: When the total pressure loss coefficient does not meet the design requirements, the air profile control parameters in the initial air profile region where the film air hole is located are adjusted individually or in combination until the total pressure loss coefficient of the new air profile meets the design requirements.

[0010] Furthermore, the set width is ±5% to 10% of the width of the throat position.

[0011] Furthermore, the load distribution of the initial airfoil and the total pressure loss coefficient of the updated airfoil are both calculated using simulation software.

[0012] Furthermore, the initial blade profile has an exit configuration angle of 70° to 140°, an installation angle of 20° to 40°, an exit airflow angle at the trailing edge of 12° to 18°, and a trailing edge bend angle of 5° to 20°.

[0013] Another aspect of this application discloses a low-consistency gas-cooled turbine guide vane design system, wherein the low-consistency gas-cooled turbine guide vane is provided with multiple exhaust film holes, and the method includes: The design module is configured to conduct preliminary design of the blade profile of the low-consistency, strongly cooled gas mixing guide vane to obtain the initial blade profile; The simulation calculation module is configured to simulate and calculate the load distribution of the initial airfoil in order to obtain the Mach number distribution curve of the suction surface of the initial airfoil. The partitioning module is configured to partition the blade channel formed by adjacent initial blade profiles, and divide the suction surface load distribution curve into multiple regions along the flow direction based on the partitioning boundary of the blade channel. The update module is configured to determine the region where the film air hole is located on the suction surface Mach number distribution curve. When at least one film air hole is located in the set region of the suction surface Mach number distribution curve, the strong cold air mixing effect of the cold air outflow from the film air hole and its influence on the air profile loss are coupled and considered. The air profile control parameters in the initial air profile region where the film air hole is located are adjusted individually or jointly to reduce the Mach number at the location of the film air hole and keep the Mach number distribution curve at other locations unchanged, so that the influence of the cold air outflow on the mixing loss of the mainstream is minimized, and a new air profile considering the influence of the cold air outflow from the film air hole is obtained. The output module is configured to calculate the total pressure loss coefficient of the new airfoil. When the total pressure loss coefficient meets the design requirements, the new airfoil is output as the design airfoil.

[0014] Compared with the prior art, this application has the following advantages: This application proposes a low-loss airfoil design method for low-consistency air-cooled turbine guide vanes based on airfoil load zoning control. This method can take into account the influence of airfoil and film film outflow at typical positions under large cooling air volume conditions, thereby significantly reducing the mixing loss of film film outflow and mainstream flow on the suction side and improving the aerodynamic performance of the turbine guide vane.

[0015] Other features and advantages of this application will be set forth in the description which follows, and will be apparent in part from the description, or may be learned by practicing the application. The objectives and other advantages of this application may be realized and obtained by means of the structures pointed out in the description, claims and drawings. Attached Figure Description

[0016] To more clearly illustrate the technical solutions in the embodiments of this application or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are some embodiments of this application. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0017] Figure 1 A comparison diagram showing the throat location and film pore location for both regular and low consistency is presented. Figure 2 The comparison of the total pressure loss coefficients of the film gas pores at different locations is shown; Figure 3 A flowchart of the method of this application is shown; Figure 4 A schematic diagram of an embodiment of the method of this application is shown; Figure 5 The diagram shows the partitioning of the initial airfoil load distribution and suction surface Mach number distribution curves; Figure 6a A diagram showing the location of the blade cascade throat is provided. Figure 6b A schematic diagram of the blade cascade channel partitioning is shown; Figure 7 The relative positions of the axial chord lengths of the air film pores on the suction surface are shown; Figure 8 A comparison diagram of load distribution between the initial airfoil and the updated airfoil is shown; Figure 9 The example shows a comparison diagram of the initial leaf shape of this application after forward optimization and backward optimization. Detailed Implementation

[0018] To make the objectives, technical solutions, and advantages of the embodiments of this application clearer, the technical solutions of the embodiments of this application will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of this application, not all embodiments. Based on the embodiments of this application, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of this application.

[0019] Existing research indicates that the mixing loss between the cold air outflow and the mainstream consists of two parts: the loss due to the temperature difference between the cold air and the mainstream, and the mixing loss due to the velocity difference between the cold air and the mainstream. Since the bleed air temperature is limited by the bleed air temperature, the temperature difference loss between the cold air and the mainstream is unavoidable. Therefore, the mixing loss due to the velocity difference between the cold air outflow and the mainstream is the main and controllable cause of the cold air outflow loss. According to... Figure 1 b. The high-load, low-consistency airfoil design places the film gas pores on the suction surface of the combustion chamber guide vanes at a higher local Mach number, inevitably leading to a sharp increase in mixing losses between the cold gas outflow and the mainstream. Simultaneously... Figure 2 The influence of the mixing losses of the cold air outflow and mainstream flow at different locations of the film cooling orifices on the aerodynamic performance (total pressure loss coefficient) of the combustion-guided turbine is presented. The total pressure loss coefficient of the film cooling orifices at different locations exhibits the following characteristics: suction side film cooling orifices > pressure side film cooling orifices > leading edge film cooling orifices. The total pressure loss coefficient at the design point on the back of the blade (near the throat) is more than 1.5 times that at the leading edge, making it the most significant factor affecting the total pressure loss coefficient of the combustion-guided turbine. Based on this, this application proposes a low-consistency air-cooled turbine guide vane design method, wherein the guide vane has multiple film cooling orifices on the suction side of the blade, referring to... Figure 3 The methods include: S100: Conduct preliminary design of the blade profile for the low-consistency guide vane to obtain the initial blade profile; S200: Simulation calculations are used to obtain the load distribution of the initial airfoil, in order to obtain the Mach number distribution curve of the suction surface of the initial airfoil; S300: The blade channel formed by adjacent initial airfoils is divided into partitions, and the suction surface load distribution curve is divided into multiple regions along the flow direction based on the partition boundaries of the blade channel. S400: Determine the region where the film air hole is located on the suction surface Mach number distribution curve. When at least one film air hole is located in the set region of the suction surface Mach number distribution curve, the strong cold air mixing effect of the cold air outflow from the film air hole and its influence on the air profile loss are coupled and considered. The air profile control parameters in the initial air profile region where the film air hole is located are adjusted individually or jointly to reduce the Mach number at the location of the film air hole and keep the Mach number distribution curve at other locations unchanged, so that the influence of the cold air outflow on the mixing loss of the mainstream is minimized, and a new air profile considering the influence of the cold air outflow from the film air hole is obtained. S500: Calculate the total pressure loss coefficient of the new airfoil. When the total pressure loss coefficient meets the design requirements, the new airfoil will be output as the design airfoil.

[0020] Specifically, refer to Figure 4 A specific example flow of the above method is given, which includes: Initial airfoil design: Without considering the influence of cold air outflow from the film gas vent, a preliminary airfoil design for the low-consistency guide vane is carried out to obtain the initial airfoil. Simulation calculation: The load distribution of the initial airfoil is simulated to obtain the Mach number distribution curve of the suction surface of the initial airfoil; Load zoning: The blade channel formed by the initial airfoil is divided into three regions: before the throat, near the throat, and after the throat. The air film vents of the suction surface are located in the divided regions. Airfoil design considering film cooling orifices: When at least one film cooling orifice is located in the set region of the suction surface, the strong cold air mixing effect of the cold air outflow from the film cooling orifice and its influence on airfoil loss are considered in a coupled manner. The airfoil control parameters in the initial airfoil region where the film cooling orifice is located are adjusted individually or jointly to reduce the Mach number at the location of the film cooling orifice, while keeping the Mach number distribution curve of other locations as unchanged as possible. This minimizes the influence of the cold air outflow on the mixing loss of the mainstream, so as to obtain a new airfoil that considers the influence of the cold air outflow from the film cooling orifice. Update airfoil: Calculate the total pressure loss coefficient of the new airfoil. When the total pressure loss coefficient meets the design requirements, the new airfoil is output as the design airfoil.

[0021] In some embodiments, the above method further includes: When the total pressure loss coefficient does not meet the design requirements, the air profile control parameters in the initial air profile region where the film air hole is located are adjusted individually or in combination until the total pressure loss coefficient of the new air profile meets the design requirements.

[0022] In some embodiments, in step S100, based on the one-dimensional design results of the gas turbine, a preliminary design of the guide vane profile for high load and low density can be performed using, for example, the eleven-parameter method. Specifically, in some examples, the number of guide vanes is 12 to 24, the outlet configuration angle (angle with the crest line) of the initial airfoil is 70° to 140°, the installation angle of the initial airfoil is 20° to 40°, the trailing edge outlet airflow angle is 12° to 18°, and the trailing edge bending angle is 5° to 20°. At the same time, during the initial airfoil design process, the relevant parameters and control points of the suction and pressure surfaces of the initial airfoil can be adjusted to make the curvature of the suction and pressure surfaces of the initial airfoil continuous.

[0023] In some embodiments, the load distribution of the initial airfoil can be calculated using CFD software. Specifically, existing CFD software can be used to perform numerical simulations of the cascade of the initial airfoil to obtain the load distribution on the surface of the initial airfoil, which is then represented by the Mach number. For example, referring to... Figure 5 This is a load distribution diagram of the initial airfoil.

[0024] In some embodiments, when the cascade channel is divided into sections, a pre-throat region, a throat region, and a post-throat region can be formed accordingly. Specifically, the process of dividing the pre-throat region, throat region, and post-throat region is as follows: (1): Draw a perpendicular line from the tangent point of the initial blade trailing edge to the suction surface of the adjacent guide vane to obtain the following... Figure 6a The location of the throat is shown; (2): The position of the throat is offset by ±5% to 10% of the width of the throat, so as to obtain the following... Figure 6b The throat area shown in region 2; (3): The region of the cascade channel covered by the upstream side of the initial blade leading edge to the throat region is defined as the pre-throat region, i.e. Figure 6b Central Region 1; (4) The region of the cascade channel covered by the trailing edge of the initial blade shape to the throat region is defined as the post-throat region, i.e. Figure 6b Central Region 3.

[0025] After dividing the blade cascade channels into sections, refer to Figure 7 and Figure 8The suction surface Mach number distribution curve can be partitioned according to the partitioning results based on the blade cascade channel. That is, the suction surface Mach number distribution curve is divided into multiple regions along the flow direction based on the partitioning boundary of the blade cascade channel. Specifically, the regional division of the suction surface Mach number distribution curve corresponds to the division of the blade cascade channel, which also includes the pre-throat region, the throat region, and the post-throat region. At this time, in the suction surface Mach number distribution curve, the pre-throat region corresponds to the airflow acceleration and Mach number rise section; the region near the throat is the Mach number peak section, which is also the region with the highest cold air mixing loss; the post-throat region corresponds to the airflow deceleration and Mach number fall section.

[0026] In some embodiments, during the determination of the location of the film gas pores, the flow direction of the pores on the suction surface can be initially determined based on experience. Specifically, they are distributed according to the axial chord length of 0 to 1. For low-consistency blade types, the film gas pores are generally arranged at positions with a relative axial chord length of 0 to 0.7. Figure 8 At this point, the location of the air film pores may fall into several possibilities: a: The air film pores are located in the anterior region of the throat; b: The air membrane pores are located in the anterior and larynx regions; c: The air film pores are located in the anterior, larynx, and posterior larynx regions.

[0027] In some embodiments, the airfoil control parameters include the initial airfoil's mounting angle, trailing edge bend angle, inlet configuration angle, leading edge wedge angle, outlet configuration angle, and surface curvature. By adjusting the airfoil control parameters of the initial airfoil, the Mach number of at least one region in the throat region, throat region, and throat rear region can be adjusted to reduce the Mach number at the location of the film cooling orifice in the suction surface. This results in a decrease in the local Mach number of the film cooling orifice by more than 5% to 20% compared to the initial airfoil, and a decrease in total pressure loss of more than 3%. At the same time, when adjusting the control parameters, the throat area of ​​the guide vane must remain unchanged. Specifically, this mainly reduces the influence of the film cooling orifice in the throat region or throat rear region. By adjusting the airfoil control parameters, the mixing loss of the mainstream by the film cooling orifice in the throat region or throat rear region can be significantly reduced.

[0028] The following is a simple example of optimizing a certain airfoil using the method of this invention. By adjusting the shape parameters and curvature of the suction surface of a prototype airfoil, an optimized airfoil is obtained (scheme 1, forward optimization). To compare with the method of this invention, a reverse design was performed, resulting in a reverse optimization example (scheme 2, reverse optimization). The airfoil load distributions of the two examples are as follows. Figure 9As shown in Table 1, the mixing loss of the cold air outflow with the mainstream was verified using numerical simulation. Table 1 shows that, using the method of this invention, the forward optimization of the airfoil (Scheme 1) reduced the local Mach number distribution of the film gas inlet in the throat region and near the throat region by controlling the airfoil load distribution in these two regions, resulting in a 5.2% reduction in the mixing loss of the cold air with the mainstream. Conversely, using the opposite method of this invention, the reverse optimization (Scheme 2) showed an increase in the local Mach number in the throat region and near the throat region, resulting in a 2.7% increase in the mixing loss of the cold air with the mainstream. These two examples sufficiently demonstrate the effectiveness of the method provided by this invention.

[0029] Table 1 Comparison of Optimization Results

[0030] Although this application 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 of the technical features; and these modifications or substitutions do not cause the essence of the corresponding technical solutions to deviate from the spirit and scope of the technical solutions of the embodiments of this application.

Claims

1. A method for designing the airfoil profile of a low-consistency air-cooled turbine guide vane, characterized in that, The guide vane has multiple exhaust film holes on the suction surface of the vane, and the method includes: Preliminary design of the blade profile for the low-consistency guide vane was carried out to obtain the initial blade profile; The load distribution of the initial airfoil is obtained through simulation calculation, so as to obtain the Mach number distribution curve of the suction surface of the initial airfoil; The blade cascade channel formed by adjacent initial blade profiles is divided into sections, and the suction surface Mach number distribution curve is divided into multiple regions along the flow direction based on the section boundaries of the blade cascade channel. The region where the film cooling orifice is located on the Mach number distribution curve of the suction surface is determined. When at least one film cooling orifice is located in the set region of the Mach number distribution curve of the suction surface, the strong cold air mixing effect of the cold air outflow from the film cooling orifice and its influence on the airfoil loss are coupled and considered. The airfoil control parameters in the initial airfoil region where the film cooling orifice is located are adjusted individually or jointly to reduce the Mach number at the location of the film cooling orifice and keep the Mach number distribution curve at other locations unchanged, so that the influence of the cold air outflow on the mixing loss of the mainstream is minimized, and a new airfoil considering the influence of the cold air outflow from the film cooling orifice is obtained. Calculate the total pressure loss coefficient of the new airfoil. When the total pressure loss coefficient meets the design requirements, the new airfoil is output as the design airfoil.

2. The method for designing the airfoil profile of a low-consistency air-cooled turbine guide vane according to claim 1, characterized in that, The blade cascade channel formed by adjacent initial airfoils is divided into zones, and the suction surface Mach number distribution curve is divided into multiple regions along the flow direction based on the zone boundaries of the blade cascade channel, including: Draw a perpendicular line from the initial blade trailing edge tangent point to the suction surface of the adjacent blade to determine the throat position of the blade passageway; The throat position is offset according to the set width to divide the throat area in the blade channel; The upstream portion of the blade passageway from the initial airfoil leading edge to the throat region is divided into the pre-throat region; The downstream portion of the blade channel from the initial blade shape trailing edge to the throat region is divided into the post-throat region; Determine the boundary points of the pre-throat region, throat region, and post-throat region in the suction surface Mach number distribution curve, and correspondingly divide the suction surface Mach number distribution curve along the flow direction into the pre-throat region, throat region, and post-throat region.

3. The method for designing the airfoil profile of a low-consistency air-cooled turbine guide vane according to claim 1, characterized in that, The coupling consideration of the strong cold air mixing effect of the cold air outflow from the film cooling orifice and its impact on airfoil losses involves individually or jointly adjusting the airfoil control parameters within the initial airfoil region where the film cooling orifice is located. This reduces the Mach number at the location of the film cooling orifice while keeping the Mach number distribution curves at other locations unchanged, minimizing the impact of the cold air outflow on the mixing loss of the mainstream. This includes individually or jointly adjusting the initial airfoil's installation angle, trailing edge bend angle, inlet configuration angle, leading edge wedge angle, outlet configuration angle, and surface curvature. This results in a Mach number at the film cooling orifice location that is 5% to 20% or more lower than the initial airfoil, and a total pressure loss that is 3% or more lower.

4. The method for designing the airfoil profile of a low-consistency air-cooled turbine guide vane according to claim 3, characterized in that, When adjusting the airfoil control parameters, keep the throat area of ​​the guide vane unchanged.

5. The method for designing the airfoil profile of a low-consistency air-cooled turbine guide vane according to claim 1, characterized in that, It can significantly reduce the mixing loss of the mainstream by the air film pores in the throat region or the post-throat region.

6. The method for designing the airfoil profile of a low-consistency air-cooled turbine guide vane according to claim 1, characterized in that, The method further includes: When the total pressure loss coefficient does not meet the design requirements, the air profile control parameters in the initial air profile region where the film air hole is located are adjusted individually or in combination until the total pressure loss coefficient of the new air profile meets the design requirements.

7. The method for designing the airfoil profile of a low-consistency air-cooled turbine guide vane according to claim 2, characterized in that, The set width is ±5% to 10% of the width of the throat position.

8. The method for designing the airfoil profile of a low-consistency air-cooled turbine guide vane according to claim 1, characterized in that, The load distribution of the initial airfoil and the total pressure loss coefficient of the new airfoil were both calculated using simulation software.

9. The method for designing the airfoil profile of a low-consistency air-cooled turbine guide vane according to claim 1, characterized in that, The initial airfoil has an exit configuration angle of 70° to 140°, an installation angle of 20° to 40°, an exit airflow angle at the trailing edge of 12° to 18°, and a trailing edge bend angle of 5° to 20°.

10. A low-consistency air-cooled turbine guide vane design system, characterized in that, The low-consistency, high-coolness gas mixing guide vanes are provided with multiple exhaust film holes, and the system includes: The design module is configured to conduct preliminary design of the blade profile of the low-consistency, strongly cooled gas mixing guide vane to obtain the initial blade profile; The simulation calculation module is configured to simulate and calculate the load distribution of the initial airfoil in order to obtain the Mach number distribution curve of the suction surface of the initial airfoil. The partitioning module is configured to partition the blade channel formed by adjacent initial blade profiles, and divide the suction surface load distribution curve into multiple regions along the flow direction based on the partitioning boundary of the blade channel. The update module is configured to determine the region where the film air hole is located on the suction surface Mach number distribution curve. When at least one film air hole is located in the set region of the suction surface Mach number distribution curve, the strong cold air mixing effect of the cold air outflow from the film air hole and its influence on the air profile loss are coupled and considered. The air profile control parameters in the initial air profile region where the film air hole is located are adjusted individually or jointly to reduce the Mach number at the location of the film air hole and keep the Mach number distribution curve at other locations unchanged, so that the influence of the cold air outflow on the mixing loss of the mainstream is minimized, and a new air profile considering the influence of the cold air outflow from the film air hole is obtained. The output module is configured to calculate the total pressure loss coefficient of the new airfoil. When the total pressure loss coefficient meets the design requirements, the new airfoil is output as the design airfoil.

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