Turbine blade cooling structure and design method thereof

Abstract

The invention relates to the technical field of aero-engine turbine blade cooling design, and discloses a turbine blade cooling structure and a design method thereof.The design method of the turbine blade cooling structure comprises the steps that first design parameters are obtained according to the overall engine and air system requirements, and second design parameters are obtained according to the weight reduction requirement and inner cavity flow resistance of turbine blades; obtaining a second design parameter; calculating test parameter deviation, and judging whether the test parameter deviation meets requirements; if the test parameter deviation does not meet the requirement, designing a flow limiting plate in front of an inlet until the calculated test parameter deviation meets the requirement; performing three-dimensional temperature field calculation on the turbine blade cooling structure; the size of the current limiting plate is locally adjusted according to the three-dimensional temperature field result, and the final structure and size of the current limiting plate are obtained. According to the design method of the turbine blade cooling structure, three-dimensional modeling does not need to be conducted on the blade again, only design adjustment needs to be conducted on the flow limiting plate, and the design process of the turbine blade cooling structure is optimized.

Description

Technical Field

[0001] The present invention relates to the technical field of aero-engine turbine blade cooling design, and in particular to a turbine blade cooling structure and a design method thereof. Background Art

[0002] Turbine blades are the most demanding components in aircraft engines that withstand thermal and mechanical loads. The role of their cooling structure is crucial. Through effective cooling, turbine blades can maintain stable performance in high temperature environments, thereby improving the thermal efficiency and performance of the entire aircraft engine.

[0003] In the prior art, when the turbine blade cooling structure needs to adjust the front and rear flow ratio, the flow area from the front and rear cavity inlets to the tenon inner cavity must be modified, the blade must be three-dimensionally modeled again, and the three-dimensional temperature field of the blade must be calculated using engineering algorithms, which is a large workload. Summary of the invention

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

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

[0006] Step S1: obtaining a first design parameter value of the turbine blade cooling structure according to overall engine and air system requirements;

[0007] Step S2: obtaining a second design parameter value of the turbine blade cooling structure according to the weight reduction requirement and the inner cavity flow resistance of the turbine blade;

[0008] Step S3: obtaining a design parameter deviation value according to the first design parameter value and the second design parameter value, and when the design parameter deviation value does not meet the prescribed value, proceeding to step S4; when the design parameter deviation value meets the prescribed value, proceeding to step S5;

[0009] Step S4: adding a flow restriction plate at the air inlet of the turbine blade, wherein the flow restriction plate is provided with a perforation, obtaining a structural parameter value of the perforation according to the design parameter deviation value, connecting the perforation with the inlet of the turbine blade to change the inner cavity flow resistance of the turbine blade cooling structure, re-obtaining a second design parameter value of the turbine blade cooling structure according to the changed inner cavity flow resistance of the turbine blade cooling structure, and performing step S3 again;

[0010] Step S5: performing a three-dimensional temperature field calculation on the turbine blade cooling structure. When there is a high-temperature concentrated area in the three-dimensional temperature field, step S6 is performed; when there is no high-temperature concentrated area in the three-dimensional temperature field, the final structural configuration of the turbine blade cooling structure is obtained;

[0011] Step S6: When the turbine blade cooling structure includes a limiting plate, the structural parameter value of the perforation is adjusted according to the result of the three-dimensional temperature field, and the second design parameter value of the turbine blade cooling structure is obtained again according to the changed inner cavity flow resistance of the turbine blade cooling structure, and step S3 is repeated; when the turbine blade cooling structure does not include a limiting plate, a limiting plate is added at the inlet of the turbine blade, and the structural parameter value of the perforation is obtained according to the result of the three-dimensional temperature field, and the second design parameter value of the turbine blade cooling structure is obtained again according to the changed inner cavity flow resistance of the turbine blade cooling structure, and step S3 is repeated.

[0012] Beneficial effect: With the design method of the turbine blade cooling structure of this embodiment, there is no need to re-model the blade in three dimensions, and only the limiting plate 30 needs to be designed and adjusted, thereby optimizing the design process of the turbine blade cooling structure.

[0013] In an optional implementation, in step S4, the structural parameter value of the perforation is obtained by the following formula:

[0014]

[0015] R`=P1*-P2;

[0016] in, is the cooling air mass flow rate, C d is the air intake flow coefficient, A is the air intake area of ​​the perforation, k is the absolute index, T1* is the total temperature, P1* is the inlet total pressure of the cold air channel, P2 is the outlet static pressure of the cold air channel, R is the gas constant, and R` is the flow resistance.

[0017] In an optional embodiment, the inner cavity flow resistance of the turbine blade includes the side inlet flow resistance of the cold air channel, the inlet flow resistance of the cold air channel, the outlet flow resistance of the cold air channel, the inner flow resistance of the cold air channel and the tail split flow resistance.

[0018] In an optional embodiment, the first design parameter value of the turbine blade cooling structure includes the actual air supply pressure of the air system and the expected flow ratio at the inlet of the cold air channel, 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 distribution ratio at the inlet of the cold air channel, and the design parameter deviation value complies with the specified numerical value that the deviation value between the required flow distribution ratio at the inlet of the cold air channel and the expected flow ratio at the inlet of the cold air channel 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 an optional embodiment, two cooling air channels are formed in the turbine blade cooling structure, and the expected inlet flow rates of the two cooling air channels are m and m, respectively. 1y and m 2y The required flow distribution ratio at the inlet of the cold air channel is m 1y / m 2y The required flow rates at the inlets of the two cold air channels are m 1x and m 2x The expected flow ratio of the cold air channel inlet is m 1x / m 2x The deviation between the required flow distribution ratio at the inlet of the cold air channel and the expected flow ratio at the inlet of the cold air channel is G, satisfying G = |m 1x / m 2x -m 1y / m 2y |×100% / (m 1y / m 2y ).

[0020] In a second aspect, the present invention provides a turbine blade cooling structure obtained using the above-mentioned design method, comprising: a blade body, an inner cavity being formed in the blade body; a partition, the partition being arranged in the inner cavity to separate the inner cavity to form at least two cold air channels, the cold air channels having an inlet and an outlet; a limiting plate, the limiting plate being arranged on the blade body and on a side of the inlet away from the outlet, at least two limiting plates being arranged, at least two limiting plates being arranged in a one-to-one correspondence with at least two cold air channels, the limiting plates being provided with through holes, and the through holes being connected to the cold air channels.

[0021] Beneficial effect: By setting the flow limiting plate, when it is necessary to adjust the flow distribution of the front and rear cavities and the flow distribution of the front and rear cavities of the blade, there is no need to re-model the blade in three dimensions. Only the flow limiting plate needs to be adjusted, which greatly reduces the workload and facilitates the design adjustment of the cooling structure.

[0022] In an optional implementation, along the extension direction of the perforation, the thickness of the limiting plate is D, satisfying 0.3 mm ≤ D ≤ 0.6 mm.

[0023] Beneficial effect: By limiting the thickness of the flow restriction plate, the structural strength of the flow restriction plate is ensured while preventing the flow restriction plate from being too thick and excessively increasing the overall weight of the engine.

[0024] In an optional embodiment, perpendicular to the extension direction of the through-hole, the cross-sectional shape of the through-hole is at least one of an ellipse, a circle and a polygon.

[0025] In an optional embodiment, a plurality of perforations are provided, and the plurality of perforations are spaced apart on the limiting plate, and the plurality of perforations are connected to the cold air channel; the cross-sectional shape of the perforations is circular, and the hole margin distance between two adjacent perforations is E, satisfying E>0.5mm.

[0026] Beneficial effect: By limiting the hole margin distance between two adjacent perforations, it is prevented that the structural strength between the two perforations is low and fracture occurs.

[0027] In an optional embodiment, a tail splitting slot is opened on the blade body, and a plurality of the tail splitting slots are provided, and the plurality of the tail splitting slots are arranged at intervals, and the plurality of the tail splitting slots are connected to a cooling air channel close to the trailing edge of the blade body.

[0028] Beneficial effect: The tail split seam is connected to the cold air channel, and the discharge of cold air through the tail split seam can ensure the smooth flow of the cooling channel inside the blade, thereby maintaining the effectiveness of the cooling system. BRIEF DESCRIPTION OF THE DRAWINGS

[0029] In order to more clearly illustrate the specific implementation methods of the present invention or the technical solutions in the prior art, the drawings required for use in the specific implementation methods or the description of the prior art will be briefly introduced below. Obviously, the drawings described below are some implementation methods of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying creative work.

[0030] Figure 1 Schematic diagram of a flow chart of a method for designing a turbine blade cooling structure according to an embodiment of the present invention.

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

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

[0033] Description of reference numerals:

[0034] 10. Blade body; 11. Inner cavity; 112. Cold air channel; 1121. Inlet; 1122. Outlet; 12. Tail split; 20. Partition; 30. Limiting plate; 31. Perforation. DETAILED DESCRIPTION

[0035] In order to make the purpose, technical solution and advantages of the embodiments of the present invention clearer, the technical solution in the embodiments of the present invention will be clearly and completely described below in conjunction with the drawings in the embodiments of the present invention. Obviously, the described embodiments are part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative work are within the scope of protection of the present invention.

[0036] Combine the following Figures 1 to 3 , describing an embodiment of the present invention.

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

[0038] Step S1: obtaining a first design parameter value of a turbine blade cooling structure according to overall engine and air system requirements;

[0039] Step S2: obtaining a second design parameter value of the turbine blade cooling structure according to the weight reduction requirement of the turbine blade and the flow resistance of the inner cavity 11;

[0040] Step S3: obtaining a design parameter deviation value according to the first design parameter value and the second design parameter value, and when the design parameter deviation value does not meet the specified value, proceeding to step S4; when the design parameter deviation value meets the specified value, proceeding to step S5;

[0041] Step S4: adding a flow restriction plate 30 at the air inlet of the turbine blade, the flow restriction plate 30 is provided with a through hole 31, obtaining a structural parameter value of the through hole 31 according to the design parameter deviation value, connecting the through hole 31 with the inlet 1121 of the turbine blade to change the flow resistance of the inner cavity 11 of the turbine blade cooling structure, re-obtaining a second design parameter value of the turbine blade cooling structure according to the changed flow resistance of the inner cavity 11 of the turbine blade cooling structure, and re-performing step S3;

[0042] Step S5: Calculate the three-dimensional temperature field of the turbine blade cooling structure. When there is a high-temperature concentrated area in the three-dimensional temperature field, proceed to step S6; when there is no high-temperature concentrated 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 limiting plate 30, the structural parameter value of the through hole 31 is adjusted according to the result of the three-dimensional temperature field, and the second design parameter value of the turbine blade cooling structure is obtained based on the changed flow resistance of the inner cavity 11 of the turbine blade cooling structure, and step S3 is repeated; when the turbine blade cooling structure does not include a limiting plate 30, a limiting plate 30 is added at the inlet 1121 of the turbine blade, and the structural parameter value of the through hole 31 is obtained according to the result of the three-dimensional temperature field, and the second design parameter value of the turbine blade cooling structure is obtained based on the changed flow resistance of the inner cavity 11 of the turbine blade cooling structure, and step S3 is repeated.

[0044] It is worth noting that, through the design method of the turbine blade cooling structure of this embodiment, there is no need to re-model the blade in three dimensions, and only the limiting plate 30 needs to be designed and adjusted, thereby 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 surface and internal temperature distribution of the turbine blade when it is in working condition. It reflects the temperature difference of the blade at different positions and depths. It is an important parameter for evaluating the blade thermal load, cooling effect and predicting the blade life.

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

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

[0048]

[0049] R`=P1*-P2;

[0050] in, is the cooling air mass flow rate, in kg / s, C d is the air intake flow coefficient, unit is dimensionless, A is the air intake area of ​​the perforation 31, unit is m 2 , k is an absolute index with a dimensionless unit, T1* is the total temperature with a unit of K, P1* is the total inlet pressure of the cold air channel 112 with a unit of Pa, P2 is the outlet static pressure of the cold air channel 112 with a unit of Pa, R is the gas constant with a dimensionless unit, and R` is the flow resistance with a unit of Pa.

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

[0052] In one embodiment, a first design parameter value of the turbine blade cooling structure includes an actual air supply pressure of the air system and an expected flow ratio at the inlet 1121 of the cold air channel 112, and a second design parameter value of the turbine blade cooling structure includes a required air supply pressure of the turbine blade and a required flow distribution ratio at the inlet of the cold air channel 112, and a design parameter deviation value meets a prescribed value that the deviation value between the required flow distribution ratio at the inlet of the cold air channel 112 and the expected flow ratio at the inlet of the cold air channel 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 the expected total cooling air flow within the turbine blade cooling structure can also be determined based on the overall engine and air system requirements, combined with the turbine blade material properties and design experience.

[0054] It should be further explained 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%, indicating that the flow resistance of the inner cavity 11 of the turbine blade is too large and the flow resistance of the inner cavity 11 needs to be reduced. y -P x ) / P y >5%, which means that under the expected total cooling air flow rate in a given turbine blade cooling structure, the blade air supply pressure is relatively small. Under the actual air supply conditions of the air system, the total cooling air flow rate to the blade will be too large, thereby affecting the engine performance. It is necessary to increase the flow resistance of the inner cavity 11 to meet the air supply pressure requirements of the air system for the blade.

[0055] In one embodiment, Figure 2 As shown, two cooling air channels 112 are formed in the turbine blade cooling structure, and the expected inlet flow rates of the two cooling air channels 112 are m and m, respectively. 1y and m 2y , the required flow distribution ratio at the inlet of the cold air channel 112 is m 1y / m 2y The required flow rates of the two cold air channels 112 are m 1x and m 2x , the expected flow rate ratio of the inlet of the cold air channel 112 is m 1x / m 2x , the deviation value between the required flow distribution ratio at the inlet of the cold air channel 112 and the expected flow ratio at the inlet of the cold air channel 112 is G, satisfying 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 at the inlet of the two cold air channels 112 is too different from the expected one, which will cause the actual cooling effect of the turbine blades to be too different from the expected one, and more high-temperature areas will appear. It is necessary to adjust the shape and size of the perforations 31 of the flow limiting plates 30 at the inlet of the two cold air channels 112 to meet the required flow distribution requirements at the inlet of the two cold air channels 112.

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

[0058] Specifically, a one-dimensional network calculation is performed on the side inlet flow resistance of the cold air channel 112, the inlet flow resistance of the cold air channel 112, the outlet flow resistance of the cold air channel 112, the flow resistance in the cavity of the cold air channel 112 and the flow resistance of the tail split slot 12 to determine the required air supply pressure of the turbine blades and the required flow distribution ratio at the inlet of the cold air channel 112.

[0059] It should be noted that one-dimensional network calculation refers to processing and analyzing one-dimensional data of the side inlet flow resistance of the cold air channel 112, the inlet flow resistance of the cold air channel 112, the outlet flow resistance of the cold air channel 112, the flow resistance in the cavity of the cold air channel 112 and the flow resistance of the tail split seam 12, and performing calculations in combination with relevant principles and formulas of fluid mechanics.

[0060] It should be noted that, in the related art, when calculating the flow resistance of the inner cavity 11, the flow resistance of the cold air flowing into the cold air channel 112 from the side intake of the cold air channel 112 is often ignored, which causes a large deviation between the calculated result and the actual experimental result; therefore, in the present embodiment, the side intake flow resistance of the cold air channel 112 is taken into account in the calculation of the flow resistance of the inner cavity 11, so as to reduce the deviation between the calculated result and the actual experimental result, reduce the design iterations of the turbine blade cooling structure, and reduce the design workload.

[0061] According to an embodiment of the present invention, in a second aspect, a turbine blade cooling structure is provided, which is obtained using the above-mentioned design method. The turbine blade cooling structure includes a blade body 10, a partition 20 and a flow limiting plate 30; an inner cavity 11 is formed in the blade body 10; the partition 20 is arranged in the inner cavity 11 to separate the inner cavity 11 to form at least two cold air channels 112, and the cold air channel 112 has an inlet 1121 and an outlet 1122; the flow limiting plate 30 is arranged on the blade body 10, and is arranged on the side of the inlet 1121 away from the outlet 1122, and at least two flow limiting plates 30 are arranged, and at least two flow limiting plates 30 are arranged one-to-one with at least two cold air channels 112, and the flow limiting plate 30 is opened with a through hole 31, and the through hole 31 is connected to the cold air channel 112.

[0062] By applying the turbine blade cooling structure of the present embodiment and setting the flow limiting plate 30, when it is necessary to adjust the flow distribution between the front and rear cavities and the front and rear cavities of the turbine blades, there is no need to re-model the turbine blades 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 the related art, when the turbine blade cooling structure needs to adjust the front-rear flow ratio, it is necessary to modify the flow area from the front and rear cavity inlets 1121 to the tenon inner cavity 11, re-model the blade in three dimensions and use engineering algorithms to calculate the three-dimensional temperature field of the blade, which is a lot of work. In this embodiment, there is no need to change the structure of the blade body 10 and the partition 20, only the flow limiting plate 30 is adjusted, thereby eliminating the three-dimensional modeling calculation of the blade body 10 and the partition 20, reducing the workload.

[0064] Specifically, the turbine blade includes a blade body 10 .

[0065] Specifically, in this embodiment, if Figure 2 As shown, the partition 20 separates the inner cavity 11 to form two cold air channels 112 , and two flow limiting plates 30 are correspondingly provided.

[0066] For details, please refer to Figure 2 The structural configurations of the two cold air channels 112 are a straight channel configuration and a winding channel configuration respectively. The cold air channel 112 with a winding channel configuration is divided into three cavities to improve the cooling efficiency, while reducing fuel consumption and extending the life of the engine. The cold air channel 112 with a straight channel configuration can avoid excessive temperature rise of the cold air in the winding channel.

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

[0068] In one embodiment, Figure 2 As shown, along the extension direction of the through hole 31, the thickness of the limiting plate 30 is D, which satisfies 0.3 mm≤D≤0.6 mm.

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

[0070] Furthermore, the material selection principle of the flow limiting plate 30 is mature processing technology, good weldability with the blade body 10 and high structural strength, and it is not easy to crack.

[0071] It should be noted that when the thickness D of the limiting plate 30 is less than 0.3 mm, the structural strength of the limiting plate 30 is low and cracks are prone to occur; when the thickness D of the limiting plate 30 is greater than 0.6 mm, the limiting plate 30 is too thick and will increase the overall weight of the engine; it is understandable that the lighter the weight of the aircraft engine, the less energy is consumed during takeoff and landing of the aircraft. Therefore, on the basis of satisfying the function of the limiting plate 30, it is necessary to reduce the weight of the limiting plate 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 limiting plate 30 as needed.

[0073] It is worth noting that by limiting the thickness of the restrictor plate 30 , the structural strength of the restrictor plate 30 is ensured while preventing the restrictor plate 30 from being too thick and excessively increasing the overall weight of the engine.

[0074] In one embodiment, perpendicular to the extension direction of the through hole 31 , the cross-sectional shape of the through hole 31 is at least one of an ellipse, a circle and a polygon.

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

[0076] It should be noted that technicians in this field can adjust the cross-sectional shape of the perforation 31 as needed. For example, when the circular diameter is larger than the width of the inlet 1121, the shape of the air inlet of the limiting plate 30 is changed to a square to meet the cross-sectional area requirements of the turbine blades for the inlet 1121 of the limiting plate 30.

[0077] In one embodiment, Figure 3 As shown, there are a plurality of perforations 31, which are spaced apart on the limiting plate 30, and are all connected to the cold air passage 112; the cross-sectional shape of the perforations 31 is circular, and the hole margin distance between two adjacent perforations 31 is E, satisfying E>0.5mm.

[0078] Specifically, Figure 3 As shown, in this embodiment, two through holes 31 are provided, and the hole margin E between two adjacent through holes 31 is 0.6 mm.

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

[0080] It should be noted that when the cold air flows into the cold air channel 112, the cold air generates pressure on the surface of the limiting plate 30. If the hole edge distance between two adjacent perforations 31 is small, the limiting plate 30 structure between the two adjacent perforations 31 may produce cracks 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 hole margin distance between two adjacent through holes 31 , it is prevented that the structural strength between the two through holes 31 is low and fracture occurs.

[0082] In one embodiment, a tail split slot 12 is opened on the blade body 10 . A plurality of tail split slots 12 are provided. The plurality of tail split slots 12 are arranged at intervals. The plurality of tail split slots 12 are connected to a cold air passage 112 close to the trailing edge of the blade body 10 .

[0083] Specifically, Figure 2 As shown, eight tail split slots 12 are provided, and the eight tail split slots 12 are all connected to the cold air channel 112 close to the trailing edge of the blade body 10 .

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

[0085] It is worth noting that the tail split slot 12 is connected to the cold air passage 112 , and the cold air is discharged to the external environment through the tail split slot 12 , which can ensure the smooth flow of the cold air passage 112 inside the blade body 10 , thereby maintaining the effectiveness of the cooling system.

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

Claims

1. A method for designing a turbine blade cooling structure, characterized in that: The design method comprises the steps of: Step S1: obtaining a first design parameter value of the turbine blade cooling structure according to overall engine and air system requirements; Step S2: obtaining a second design parameter value of the cooling structure of the turbine blade according to the weight reduction requirement of the turbine blade and the flow resistance of the inner cavity (11); Step S3: obtaining a design parameter deviation value according to the first design parameter value and the second design parameter value, and when the design parameter deviation value does not meet the prescribed value, proceeding to step S4; when the design parameter deviation value meets the prescribed value, proceeding to step S5; Step S4: adding a flow restriction plate (30) at the air inlet of the turbine blade, the flow restriction plate (30) being provided with a perforation (31), obtaining a structural parameter value of the perforation (31) according to the design parameter deviation value, connecting the perforation (31) with the inlet (1121) of the turbine blade to change the inner cavity flow resistance of the turbine blade cooling structure, re-obtaining a second design parameter value of the turbine blade cooling structure according to the changed inner cavity flow resistance of the turbine blade cooling structure, and performing step S3 again; Step S5: performing a three-dimensional temperature field calculation on the turbine blade cooling structure. When there is a high-temperature concentrated area in the three-dimensional temperature field, step S6 is performed; when there is no high-temperature concentrated area in the three-dimensional temperature field, the final structural configuration of the turbine blade cooling structure is obtained; Step S6: When the turbine blade cooling structure includes a limiting plate (30), the structural parameter value of the perforation (31) is adjusted according to the result of the three-dimensional temperature field, and the second design parameter value of the turbine blade cooling structure is obtained again according to the changed inner cavity flow resistance of the turbine blade cooling structure, and step S3 is repeated; when the turbine blade cooling structure does not include a limiting plate (30), a limiting plate (30) is added at the inlet (1121) of the turbine blade, and the structural parameter value of the perforation (31) is obtained according to the result of the three-dimensional temperature field, and the second design parameter value of the turbine blade cooling structure is obtained again according to the changed inner cavity flow resistance of the turbine blade cooling structure, and step S3 is repeated.

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

3. The method for designing a turbine blade cooling structure according to claim 1, characterized in that: The inner cavity flow resistance of the turbine blade includes the side inlet flow resistance of the cold air channel, the inlet flow resistance of the cold air channel, the outlet flow resistance of the cold air channel, the inner flow resistance of the cold air channel and the tail split flow resistance.

4. The method for designing a turbine blade cooling structure according to claim 1, characterized in that: The first design parameter value of the turbine blade cooling structure includes the actual air supply pressure of the air system and the expected flow ratio at the inlet of the cold air channel (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 distribution ratio at the inlet of the cold air channel (112). The design parameter deviation value meets the specified value that the deviation value between the required flow distribution ratio at the inlet of the cold air channel (112) and the expected flow ratio at the inlet of the cold air channel (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%.

5. The method for designing a turbine blade cooling structure according to claim 4, characterized in that: Two cold air passages (112) are formed in the turbine blade cooling structure, and the expected inlet flow rates of the two cold air passages (112) are m 1y and m 2y The required flow distribution ratio at the inlet of the cold air channel (112) is m 1y / m 2y The required flow rates at the inlets of the two cold air passages (112) are m 1x and m 2x The expected flow rate ratio of the inlet of the cold air channel (112) is m 1x / m 2x The deviation value between the required flow distribution ratio at the inlet of the cold air channel (112) and the expected flow ratio at the inlet of the cold air channel (112) is G, and G=|m 1x / m 2x -m 1y / m 2y |×100% / (m 1y / m 2y ).

6. A turbine blade cooling structure, obtained by using the design method according to any one of claims 1 to 5, characterized in that: include: A blade body (10), wherein an inner cavity (11) is formed in the blade body (10); a partition (20), the partition (20) being arranged in the inner cavity (11) so as to partition the inner cavity (11) to form at least two cold air passages (112), the cold air passage (112) having an inlet (1121) and an outlet (1122); A flow limiting plate (30), the flow limiting plate (30) being arranged on the blade body (10) and arranged on a side of the inlet (1121) away from the outlet (1122), at least two flow limiting plates (30) being arranged, at least two flow limiting plates (30) being arranged in a one-to-one correspondence with at least two cold air passages (112), and the flow limiting plate (30) being provided with a through hole (31), and the through hole (31) being in communication with the cold air passage (112).

7. The turbine blade cooling structure according to claim 6, characterized in that: Along the extension direction of the perforation (31), the thickness of the flow limiting plate (30) is D, satisfying 0.3 mm ≤ D ≤ 0.6 mm.

8. The turbine blade cooling structure according to claim 6, characterized in that: Perpendicular to the extension direction of the through hole (31), the cross-sectional shape of the through hole (31) is at least one of an ellipse, a circle and a polygon.

9. The turbine blade cooling structure according to claim 8, characterized in that: A plurality of the perforations (31) are provided, and the plurality of the perforations (31) are arranged at intervals on the flow limiting plate (30), and the plurality of the perforations (31) are all connected to the cold air channel (112); the cross-sectional shape of the perforations (31) is circular, and the hole margin distance between two adjacent perforations (31) is E, satisfying E>0.5 mm.

10. The turbine blade cooling structure according to claim 6, characterized in that: The blade body (10) is provided with a tail splitting slot (12), a plurality of the tail splitting slots (12) are provided, the plurality of the tail splitting slots (12) are arranged at intervals, and the plurality of the tail splitting slots (12) are connected to a cold air passage (112) close to the trailing edge of the blade body (10).

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