Turbine blade cooling structure and design method thereof

Abstract

This invention relates to the field of turbine blade cooling design technology, and discloses a turbine blade cooling structure and its design method. The design method of the turbine blade cooling structure includes the following steps: obtaining first design parameters based on the overall engine and air system requirements; obtaining second design parameters based on the turbine blade's weight reduction requirements and internal flow resistance; calculating the deviation of the test parameters and determining whether they meet the requirements; if they do not meet the requirements, designing a flow restrictor before the inlet until the calculated deviation of the test parameters meets the requirements; performing a three-dimensional temperature field calculation on the turbine blade cooling structure; and locally adjusting the size of the flow restrictor based on the three-dimensional temperature field results to obtain the final structure and size of the flow restrictor. This invention, through the turbine blade cooling structure design method of this embodiment, eliminates the need for re-modeling the blade in three dimensions, requiring only design adjustments to the flow restrictor, thus optimizing the design process of the turbine blade cooling structure.

Description

Technical Field

[0001] This invention relates to the field of turbine blade cooling design technology, specifically to turbine blade cooling structures and their design methods. Background Technology

[0002] Turbine blades are the most demanding components in aero engines in terms of thermal and mechanical loads, and their cooling structure plays a crucial role. Effective cooling can maintain stable performance of turbine blades in high-temperature environments, thereby improving the thermal efficiency and performance of the entire aero engine.

[0003] In existing technologies, when the cooling structure of turbine blades needs to adjust the front and rear flow ratio, the flow area from the front and rear cavity inlet to the tenon cavity must be modified, the blade must be remodeled in three dimensions, and the three-dimensional temperature field of the blade must be calculated using engineering algorithms, which involves a large amount of work. Summary of the Invention

[0004] In view of this, the present invention provides a turbine blade cooling structure and its design method to solve the problem of large workload when adjusting the front and rear flow ratio of the turbine blade cooling structure.

[0005] In a first aspect, the present invention also provides a design method for a turbine blade cooling structure, the design method comprising the steps of:

[0006] Step S1: Obtain the first design parameter values ​​of the turbine blade cooling structure based on the overall engine and air system requirements;

[0007] Step S2: Based on the weight reduction requirements of the turbine blades and the internal flow resistance, obtain the second design parameter values ​​of the turbine blade cooling structure;

[0008] Step S3: Obtain the design parameter deviation value based on the first design parameter value and the second design parameter value. If the design parameter deviation value does not meet the specified value, proceed to step S4; if the design parameter deviation value meets the specified value, proceed to step S5.

[0009] Step S4: Add a flow restrictor at the inlet of the turbine blade. The flow restrictor has a perforation. Based on the design parameter deviation value, obtain the structural parameter value of the perforation. Connect the perforation to the inlet of the turbine blade to change the internal flow resistance of the turbine blade cooling structure. Based on the changed internal flow resistance of the turbine blade cooling structure, obtain the second design parameter value of the turbine blade cooling structure again, and repeat step S3.

[0010] Step S5: Perform three-dimensional temperature field calculation on the turbine blade cooling structure. If there is a high-temperature concentration area in the three-dimensional temperature field, proceed to step S6; if there is no high-temperature concentration area in the three-dimensional temperature field, obtain the final structural configuration of the turbine blade cooling structure.

[0011] Step S6: When the turbine blade cooling structure includes a flow restrictor, the structural parameter values ​​of the perforation are adjusted according to the results of the three-dimensional temperature field. The second design parameter values ​​of the turbine blade cooling structure are obtained again according to the changed internal flow resistance of the turbine blade cooling structure, and step S3 is repeated. When the turbine blade cooling structure does not include a flow restrictor, a flow restrictor is added at the inlet of the turbine blade. The structural parameter values ​​of the perforation are obtained according to the results of the three-dimensional temperature field. The second design parameter values ​​of the turbine blade cooling structure are obtained again according to the changed internal flow resistance of the turbine blade cooling structure, and step S3 is repeated.

[0012] Beneficial effects: The design method of the turbine blade cooling structure in this embodiment eliminates the need to remodel the blade in three dimensions, and only requires design adjustments to the flow restrictor 30, thus optimizing the design process of the turbine blade cooling structure.

[0013] In an optional implementation, in step S4, the structural parameter values ​​of the perforation are obtained using the following formula:

[0014]

[0015] R` = P1*-P2;

[0016] in, C is the mass flow rate of the cooling air. d denoted as , where A is the intake flow rate coefficient, k is the absolute index, T1* is the total temperature, P1* is the total inlet pressure of the cold air passage, P2 is the static pressure at the outlet of the cold air passage, R is the gas constant, and R' is the flow resistance.

[0017] In one optional embodiment, the internal flow resistance of the turbine blade includes the flow resistance of the cold air passage side inlet, the flow resistance of the cold air passage inlet, the flow resistance of the cold air passage outlet, the flow resistance inside the cold air passage cavity, and the flow resistance of the tail slot.

[0018] In one optional embodiment, the first design parameter value of the turbine blade cooling structure includes the ratio of the actual air supply pressure of the air system to the expected inlet flow rate of the cold air passage, and the second design parameter value of the turbine blade cooling structure includes the required air supply pressure of the turbine blade and the required inlet flow rate distribution ratio of the cold air passage. The deviation value of the design parameter meets the specified value that the deviation between the required inlet flow rate distribution ratio of the cold air passage and the expected inlet flow rate ratio of the cold air passage is less than 15%, and the deviation between the required air supply pressure of the blade and the actual air supply pressure of the air system is less than 5%.

[0019] In one optional embodiment, two cooling air channels are formed within the turbine blade cooling structure, and the expected inlet flow rates of the two cooling air channels are respectively m 1y and m 2y The required flow distribution ratio at the inlet of the cold air passage is m. 1y / m 2y The required inlet flow rates for the two aforementioned cold air passages are respectively m 1x and m 2x The expected inlet flow rate ratio of the cold air passage is m 1x / m 2x The deviation between the required flow rate allocation ratio at the inlet of the cold air passage and the expected flow rate ratio at the inlet of the cold air passage is G, satisfying G = |m 1x / m 2x -m 1y / m 2y |×100% / (m 1y / m 2y ).

[0020] Secondly, the present invention provides a turbine blade cooling structure obtained using the above-described design method, comprising: a blade body having an inner cavity formed therein; a baffle plate disposed within the inner cavity to divide the inner cavity into at least two cooling air passages, the cooling air passages having an inlet and an outlet; and a flow restrictor plate disposed on the blade body and on the side of the inlet away from the outlet, wherein at least two flow restrictors are provided, and the at least two flow restrictors are correspondingly disposed to at least two cooling air passages, and the flow restrictor plate has a perforation that communicates with the cooling air passage.

[0021] Beneficial effects: By setting up a flow restrictor, when it is necessary to adjust the flow distribution between the front and rear cavities and the flow distribution between the front and rear cavities of the blades, there is no need to remodel the blades in three dimensions. Only the flow restrictor needs to be adjusted, which greatly reduces the workload and facilitates the design and adjustment of the cooling structure.

[0022] In one alternative embodiment, the thickness of the flow restrictor is D along the extension direction of the perforation, satisfying 0.3mm≤D≤0.6mm.

[0023] Beneficial effects: By limiting the thickness of the flow restrictor, the structural strength of the flow restrictor is ensured while avoiding excessive increase in the overall weight of the engine due to excessive thickness.

[0024] In one alternative embodiment, the cross-sectional shape of the perforation is at least one of elliptical, circular, and polygonal, perpendicular to the extending direction of the perforation.

[0025] In one optional embodiment, a plurality of perforations are provided, and the plurality of perforations are spaced apart on the flow-limiting plate, and all of the plurality of perforations are connected to the cold air channel; the cross-sectional shape of the perforation is circular, and the distance between the edges of two adjacent perforations is E, which satisfies E > 0.5 mm.

[0026] Beneficial effect: By limiting the distance between the edges of two adjacent perforations, the structural strength between the two perforations is low, preventing breakage.

[0027] In one optional embodiment, a tail slit is provided on the blade body, and a plurality of tail slits are provided, the plurality of tail slits being spaced apart, and the plurality of tail slits being connected to a cooling air channel near the tail edge of the blade body.

[0028] Beneficial effects: The tail slit is connected to the cooling air passage, and the cooling air is discharged through the tail slit, which can ensure the smooth flow of the internal cooling passage of the blade, thereby maintaining the effectiveness of the cooling system. Attached Figure Description

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

[0030] Figure 1 This is a schematic flowchart illustrating the design method of the turbine blade cooling structure according to an embodiment of the present invention.

[0031] Figure 2 This is a schematic diagram of the turbine blade cooling structure according to an embodiment of the present invention;

[0032] Figure 3 This is a schematic diagram of the current limiting plate according to an embodiment of the present invention;

[0033] Explanation of reference numerals in the attached figures:

[0034] 10. Blade; 11. Inner cavity; 112. Cold air passage; 1121. Inlet; 1122. Outlet; 12. Tail split; 20. Baffle; 30. Flow restrictor; 31. Perforation. Detailed Implementation

[0035] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0036] The following is combined with Figures 1 to 3 The following describes embodiments of the present invention.

[0037] According to an embodiment of the present invention, in a first aspect, such as Figure 1 As shown, a design method for a turbine blade cooling structure is provided, which includes the following steps:

[0038] Step S1: Obtain the first design parameter values ​​for the turbine blade cooling structure based on the overall engine and air system requirements;

[0039] Step S2: Based on the weight reduction requirements of the turbine blades and the flow resistance of the inner cavity 11, obtain the second design parameter values ​​of the turbine blade cooling structure;

[0040] Step S3: Obtain the design parameter deviation value based on the first design parameter value and the second design parameter value. If the design parameter deviation value does not meet the specified value, proceed to step S4; if the design parameter deviation value meets the specified value, proceed to step S5.

[0041] Step S4: Add a flow restrictor 30 at the inlet of the turbine blade. The flow restrictor 30 has a perforation 31. According to the design parameter deviation value, obtain the structural parameter value of the perforation 31. Connect the perforation 31 to the inlet 1121 of the turbine blade to change the flow resistance of the inner cavity 11 of the turbine blade cooling structure. According to the changed flow resistance of the inner cavity 11 of the turbine blade cooling structure, obtain the second design parameter value of the turbine blade cooling structure again, and repeat step S3.

[0042] Step S5: Perform three-dimensional temperature field calculation on the turbine blade cooling structure. If there is a high-temperature concentration area in the three-dimensional temperature field, proceed to step S6; if there is no high-temperature concentration area in the three-dimensional temperature field, obtain the final structural configuration of the turbine blade cooling structure.

[0043] Step S6: When the turbine blade cooling structure includes a flow restrictor 30, the structural parameter values ​​of the perforation 31 are adjusted according to the results of the three-dimensional temperature field. The second design parameter values ​​of the turbine blade cooling structure are obtained again according to the flow resistance of the inner cavity 11 of the changed turbine blade cooling structure, and step S3 is repeated. When the turbine blade cooling structure does not include a flow restrictor 30, a flow restrictor 30 is added at the inlet 1121 of the turbine blade. The structural parameter values ​​of the perforation 31 are obtained according to the results of the three-dimensional temperature field. The second design parameter values ​​of the turbine blade cooling structure are obtained again according to the flow resistance of the inner cavity 11 of the changed turbine blade cooling structure, and step S3 is repeated.

[0044] It is worth noting that the design method of the turbine blade cooling structure in this embodiment does not require re-modeling the blade in three dimensions, but only requires designing and adjusting the flow restrictor 30, thus optimizing the design process of the turbine blade cooling structure.

[0045] It should be noted that the three-dimensional temperature field refers to the three-dimensional spatial state of the temperature distribution on the surface and inside of the turbine blade under working conditions. It reflects the temperature difference at different locations and depths of the blade and is an important parameter for evaluating the blade's thermal load, cooling effect, and predicting blade life.

[0046] It should be noted that flow resistance refers to the resistance encountered by cold air when passing through the cold air channel 112 under stable airflow conditions.

[0047] In one embodiment, in step S4, the structural parameter values ​​of the perforation 31 are obtained by the following formula:

[0048]

[0049] R` = P1*-P2;

[0050] in, This refers to the mass flow rate of the cooling air, expressed in kg / s, C. d The airflow coefficient is dimensionless, and A is the airflow area of ​​perforation 31, in meters (m²). 2 k is the absolute exponent, in dimensionless units; T1* is the total temperature, in K; P1* is the total inlet pressure of the cold air passage 112, in Pa; P2 is the static pressure at the outlet of the cold air passage 112, in Pa; R is the gas constant, in dimensionless units; and R` is the flow resistance, in Pa.

[0051] Specifically, the structural parameter of the perforation 31 is the flow resistance of the perforation 31.

[0052] In one embodiment, the first design parameter value of the turbine blade cooling structure includes the ratio of the actual air supply pressure of the air system to the expected flow rate of the inlet 1121 of the cooling air passage 112, and the second design parameter value of the turbine blade cooling structure includes the required air supply pressure of the turbine blade and the required flow rate distribution ratio of the inlet of the cooling air passage 112. The design parameter deviation value meets the specified value that the deviation between the required flow rate distribution ratio of the inlet of the cooling air passage 112 and the expected flow rate ratio of the inlet of the cooling air passage 112 is less than 15%, and the deviation between the required air supply pressure of the blade and the actual air supply pressure of the air system is less than 5%.

[0053] It should be noted that, based on the overall engine and air system requirements, combined with the turbine blade material properties and design experience, the expected total cooling airflow within the turbine blade cooling structure can also be determined.

[0054] It should be further noted that the pressure required for the blade air supply is P. X The actual air supply pressure of the air system is P. y When (P) x -P y ) / P y >5% indicates that the internal flow resistance of the turbine blade is too high, and the internal flow resistance needs to be reduced. When (P y -P x ) / P y A reading >5% indicates that, given the expected total cooling airflow within the turbine blade cooling structure, the pressure supplied to the blade is relatively low. Under actual air supply conditions from the air system, the total cooling airflow to the blade will be excessive, thus affecting engine performance. Therefore, it is necessary to increase the flow resistance of the internal cavity 11 to meet the air supply pressure requirements of the air system to the blade.

[0055] In one embodiment, such as Figure 2 As shown, two cold air channels 112 are formed within the turbine blade cooling structure, and the expected inlet flow rates of the two cold air channels 112 are respectively m 1y and m 2y The required flow distribution ratio at the inlet of the cold air passage 112 is m. 1y / m 2y The required inlet flow rates for the two cold air passages 112 are respectively m 1x and m 2x The expected inlet flow rate ratio of cold air passage 112 is m 1x / m 2x The deviation between the required flow rate allocation ratio at the inlet of the cold air passage 112 and the expected flow rate ratio at the inlet of the cold air passage 112 is G, which satisfies G=|m 1x / m 2x -m 1y / m 2y |×100% / (m 1y / m2y ).

[0056] It should be noted that when G > 15%, it means that the required flow distribution ratio of the inlet of the two cold air passages 112 is too different from the expectation, which will cause the actual cooling effect of the turbine blades to be too different from the expectation, resulting in more high temperature zones. It is necessary to adjust the shape and size of the perforation 31 of the flow restrictor 30 at the inlet of the two cold air passages 112 to meet the required flow distribution requirements of the inlet of the two cold air passages 112.

[0057] In one embodiment, the flow resistance of the inner cavity 11 of the turbine blade includes the inlet flow resistance on the side of the cold air passage 112, the inlet flow resistance of the cold air passage 112, the outlet flow resistance of the cold air passage 112, the flow resistance inside the cavity of the cold air passage 112, and the flow resistance of the tail slot 12.

[0058] Specifically, a one-dimensional network calculation is performed on the side airflow resistance of the cold air passage 112, the inlet flow resistance of the cold air passage 112, the outlet flow resistance of the cold air passage 112, the internal flow resistance of the cold air passage 112, and the flow resistance of the tail slit 12 to determine the required air pressure for the turbine blades and the required flow distribution ratio at the inlet of the cold air passage 112.

[0059] It should be noted that one-dimensional network calculation refers to processing and analyzing one-dimensional data on the airflow resistance on the side of the cold air passage 112, the inlet resistance of the cold air passage 112, the outlet resistance of the cold air passage 112, the internal resistance of the cold air passage 112, and the flow resistance of the tail slot 12, and then performing calculations based on relevant principles and formulas of fluid mechanics.

[0060] It should be noted that in related technologies, when calculating the flow resistance of the inner cavity 11, the flow resistance of the cold air entering the cold air channel 112 from the side is often ignored, which leads to a large deviation between the calculation results and the actual experimental results. Therefore, in this embodiment, the flow resistance of the cold air channel 112 from the side is taken into account in the calculation of the flow resistance of the inner cavity 11, thereby reducing the deviation between the calculation results and the actual experimental results, reducing the design iteration of the turbine blade cooling structure, and reducing the design workload.

[0061] According to an embodiment of the present invention, in a second aspect, a turbine blade cooling structure is provided, obtained using the above-described design method. The turbine blade cooling structure includes a blade body 10, a baffle plate 20, and a flow restrictor 30. An inner cavity 11 is formed within the blade body 10. The baffle plate 20 is disposed within the inner cavity 11 to divide the inner cavity 11 into at least two cold air passages 112, each cold air passage 112 having an inlet 1121 and an outlet 1122. The flow restrictor 30 is disposed on the blade body 10 and is disposed on the side of the inlet 1121 away from the outlet 1122. At least two flow restrictors 30 are provided, and each of the at least two flow restrictors 30 corresponds to one of the at least two cold air passages 112. The flow restrictor 30 has a perforation 31, which communicates with the cold air passage 112.

[0062] By using the turbine blade cooling structure of this embodiment and setting the flow limiting plate 30, when it is necessary to adjust the flow distribution between the front and rear chambers and the flow distribution between the front and rear chambers of the turbine blade, it is not necessary to remodel the turbine blade in three dimensions. Only the flow limiting plate 30 needs to be adjusted, which greatly reduces the workload and facilitates the design and adjustment of the cooling structure.

[0063] It should be noted that in related technologies, when the flow ratio of the front and rear sections of the turbine blade cooling structure needs to be adjusted, the flow area from the inlet 1121 of the front and rear chambers to the inner cavity 11 of the tenon must be modified, the three-dimensional model of the blade must be re-modeled, and the three-dimensional temperature field of the blade must be calculated using engineering algorithms, which involves a large amount of work. In this embodiment, there is no need to change the structure of the blade body 10 and the baffle 20, and only the flow limiting plate 30 needs to be adjusted. Therefore, the three-dimensional modeling and calculation of the blade body 10 and the baffle 20 are eliminated, reducing the workload.

[0064] Specifically, the turbine blades include the blade body 10.

[0065] Specifically, in this embodiment, such as Figure 2 As shown, the partition 20 divides the inner cavity 11 into two cold air channels 112, and two flow restrictors 30 are correspondingly provided.

[0066] For details, please refer to Figure 2 The two air cooling passages 112 have structural configurations of a straight passage and a meandering passage. The meandering passage 112 is divided into three chambers to improve cooling efficiency, reduce fuel consumption and extend engine life. The straight passage 112 can prevent the air cooling from overheating in the meandering passage.

[0067] It should be noted that those skilled in the art can adjust the specific number of the cold air passage 112 and the flow restrictor 30 as needed.

[0068] In one embodiment, such as Figure 2 As shown, along the extension direction of the perforation 31, the thickness of the flow restrictor 30 is D, which satisfies 0.3mm≤D≤0.6mm.

[0069] Specifically, in this embodiment, the thickness of the flow limiting plate 30 is 0.5 mm.

[0070] Furthermore, the material selection principle for the flow restrictor 30 is that it has mature processing technology, good weldability with the blade 10, high structural strength, and is not prone to cracking.

[0071] It should be noted that when the thickness D of the flow limiter 30 is less than 0.3 mm, the structural strength of the flow limiter 30 is low and it is prone to cracking; when the thickness D of the flow limiter 30 is greater than 0.6 mm, the thickness of the flow limiter 30 is too thick, which will increase the overall weight of the engine. It is understandable that the lighter the weight of the aircraft engine, the less energy the aircraft consumes during takeoff and landing. Therefore, while meeting the function of the flow limiter 30, it is necessary to reduce the weight of the flow limiter 30 as much as possible to avoid excessive increase in engine weight.

[0072] It should be noted that those skilled in the art can adjust the thickness of the flow restrictor 30 as needed.

[0073] It is worth noting that by limiting the thickness of the flow limiter 30, the structural strength of the flow limiter 30 is ensured while avoiding excessive increase in the overall weight of the engine due to excessive thickness.

[0074] In one embodiment, perpendicular to the extending direction of the perforation 31, the cross-sectional shape of the perforation 31 is at least one of ellipse, circle and polygon.

[0075] Specifically, such as Figure 3 As shown, in this embodiment, the cross-sectional shape of the perforation 31 is circular.

[0076] It should be noted that those skilled in the art can adjust the cross-sectional shape of the perforation 31 as needed. For example, when the circular diameter is greater than the width of the inlet 1121, the shape of the air inlet of the flow restrictor 30 can be changed to a square to meet the cross-sectional area requirements of the turbine blades for the inlet 1121 of the flow restrictor 30.

[0077] In one embodiment, such as Figure 3 As shown, there are several perforations 31, which are spaced apart on the flow restrictor 30. All perforations 31 are connected to the cold air channel 112. The cross-sectional shape of the perforation 31 is circular, and the distance between the edges of two adjacent perforations 31 is E, which satisfies E > 0.5 mm.

[0078] Specifically, such as Figure 3 As shown, in this embodiment, there are two perforations 31, and the edge distance E between two adjacent perforations 31 is 0.6mm.

[0079] It should be noted that the cross-sectional shape and size of the perforation 31 are determined based on the air intake area of ​​the two cold air channels 112 and the relative size relationship of the two inlets 1121. Those skilled in the art can adjust the number of perforations 31 and the distance between the edges of two adjacent perforations 31 as needed.

[0080] It should be noted that when cold air flows into the cold air passage 112, the cold air exerts pressure on the surface of the flow restrictor 30. If the distance between the edge of two adjacent perforations 31 is small, the flow restrictor 30 structure between the two adjacent perforations 31 may crack due to pressure, thereby affecting the cooling performance and safety performance of the turbine blade cooling structure.

[0081] It is worth noting that by limiting the distance between the edges of two adjacent perforations 31, the structural strength between the two perforations 31 is reduced, preventing breakage.

[0082] In one embodiment, a tail slit 12 is provided on the blade 10, and a plurality of tail slits 12 are provided, which are spaced apart, and the plurality of tail slits 12 are connected to a cold air channel 112 near the tail edge of the blade 10.

[0083] Specifically, such as Figure 2 As shown, there are eight tail slits 12, and all eight tail slits 12 are connected to the cold air channel 112 near the tail edge of the blade 10.

[0084] It should be noted that those skilled in the art can adjust the specific number of tail-split slots 12 as needed.

[0085] It is worth noting that the tail slit 12 is connected to the cooling air passage 112. The cooling air is discharged to the external environment through the tail slit 12, which can ensure the unobstructed flow of the cooling air passage 112 inside the blade 10, thereby maintaining the effectiveness of the cooling system.

[0086] Although embodiments of the invention have been described in conjunction with the accompanying drawings, those skilled in the art can make various modifications and variations without departing from the spirit and scope of the invention, and all such modifications and variations fall within the scope defined by the invention.

Claims

1. A design method for a turbine blade cooling structure, characterized in that, The design method includes the following steps: Step S1: Obtain the first design parameter values ​​of the turbine blade cooling structure based on the overall engine and air system requirements; Step S2: Based on the weight reduction requirements of the turbine blades and the flow resistance of the inner cavity (11), obtain the second design parameter values ​​of the turbine blade cooling structure; Step S3: Obtain the design parameter deviation value based on the first design parameter value and the second design parameter value. If the design parameter deviation value does not meet the specified value, proceed to step S4; if the design parameter deviation value meets the specified value, proceed to step S5. Step S4: Add a flow restrictor (30) at the inlet of the turbine blade. The flow restrictor (30) has a perforation (31). According to the design parameter deviation value, obtain the structural parameter value of the perforation (31). Connect the perforation (31) to the inlet (1121) of the turbine blade to change the internal flow resistance of the turbine blade cooling structure. According to the changed internal flow resistance of the turbine blade cooling structure, obtain the second design parameter value of the turbine blade cooling structure again, and repeat step S3. Step S5: Perform three-dimensional temperature field calculation on the turbine blade cooling structure. If there is a high-temperature concentration area in the three-dimensional temperature field, proceed to step S6; if there is no high-temperature concentration area in the three-dimensional temperature field, obtain the final structural configuration of the turbine blade cooling structure. Step S6: When the turbine blade cooling structure includes a flow restrictor (30), the structural parameter values ​​of the perforation (31) are adjusted according to the results of the three-dimensional temperature field. The second design parameter value of the turbine blade cooling structure is obtained again according to the internal flow resistance of the changed turbine blade cooling structure, and step S3 is repeated. When the turbine blade cooling structure does not include a flow restrictor (30), a flow restrictor (30) is added at the inlet (1121) of the turbine blade. The structural parameter values ​​of the perforation (31) are obtained according to the results of the three-dimensional temperature field. The second design parameter value of the turbine blade cooling structure is obtained again according to the internal flow resistance of the changed turbine blade cooling structure, and step S3 is repeated. The first design parameter value of the turbine blade cooling structure includes the ratio of the actual air supply pressure of the air system to the expected inlet flow rate of the cold air passage (112). The second design parameter value of the turbine blade cooling structure includes the required air supply pressure of the turbine blade and the required inlet flow rate distribution ratio of the cold air passage (112). The deviation value of the design parameter meets the specified value that the deviation between the required inlet flow rate distribution ratio of the cold air passage (112) and the expected inlet flow rate ratio of the cold air passage (112) is less than 15%, and the deviation between the required air supply pressure of the blade and the actual air supply pressure of the air system is less than 5%. The turbine blade cooling structure forms two cold air channels (112), and the expected inlet flow rates of the two cold air channels (112) are respectively m 1y and m 2y The required flow rate allocation ratio at the inlet of the cold air passage (112) is m 1y / m 2y The required inlet flow rates of the two aforementioned cold air passages (112) are respectively m 1x and m 2x The expected inlet flow rate ratio of the cold air passage (112) is m 1x / m 2x The deviation between the required flow rate allocation ratio at the inlet of the cold air passage (112) and the expected flow rate ratio at the inlet of the cold air passage (112) is G, which satisfies G=|m 1x / m 2x -m 1y / m 2y |×100% / (m 1y / m 2y ).

2. The design method for the turbine blade cooling structure according to claim 1, characterized in that, In step S4, the structural parameter values ​​of the perforation (31) are obtained by the following formula: ； R` = P1 * - P2; in, C is the mass flow rate of the cooling air. d denoted as , where A is the intake flow rate coefficient, k is the absolute index, T1* is the total temperature, P1* is the total inlet pressure of the cold air passage, P2 is the static pressure at the outlet of the cold air passage, R is the gas constant, and R' is the flow resistance.

3. The design method for the turbine blade cooling structure according to claim 1, characterized in that, The internal flow resistance of the turbine blade includes the flow resistance of the air intake on the side of the cold air passage, the flow resistance of the cold air passage inlet, the flow resistance of the cold air passage outlet, the flow resistance inside the cold air passage cavity, and the flow resistance of the tail slot.

4. A turbine blade cooling structure, obtained using the design method according to any one of claims 1 to 3, characterized in that, include: Leaf body (10), wherein an inner cavity (11) is formed within the leaf body (10); A partition (20) is disposed in the inner cavity (11) to divide the inner cavity (11) into at least two cold air passages (112), the cold air passages (112) having an inlet (1121) and an outlet (1122). A flow restrictor (30) is provided on the blade (10) and on the side of the inlet (1121) away from the outlet (1122). At least two flow restrictors (30) are provided, and at least two flow restrictors (30) are provided in correspondence with at least two cold air channels (112). The flow restrictor (30) is provided with a perforation (31), and the perforation (31) is connected to the cold air channel (112).

5. The turbine blade cooling structure according to claim 4, characterized in that, Along the extension direction of the perforation (31), the thickness of the flow restrictor (30) is D, which satisfies 0.3mm≤D≤0.6mm.

6. The turbine blade cooling structure according to claim 4, characterized in that, Perpendicular to the extending direction of the perforation (31), the cross-sectional shape of the perforation (31) is at least one of ellipse, circle and polygon.

7. The turbine blade cooling structure according to claim 6, characterized in that, The perforation (31) is provided in a plurality of places, and the plurality of perforations (31) are spaced apart on the flow limiting plate (30). The plurality of perforations (31) are all connected to the cold air channel (112). The cross-sectional shape of the perforation (31) is circular, and the distance between the edges of two adjacent perforations (31) is E, which satisfies E>0.5mm.

8. The turbine blade cooling structure according to claim 4, characterized in that, The blade (10) has a tail slit (12) provided. There are several tail slits (12) provided. The several tail slits (12) are spaced apart. The several tail slits (12) are connected to a cold air channel (112) near the tail edge of the blade (10).

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