Turbine blade air film cooling hole design method and turbine blade air film cooling hole

Abstract

The invention relates to the technical field of turbine blade cooling, in particular to a turbine blade air film cooling hole design method and a turbine blade air film cooling hole. The turbine blade air film cooling hole design method comprises the following steps that S1, modeling is conducted on a turbine blade with a high-temperature area, and related data are obtained; s2, according to the model and the data obtained in the step S1, calculating data of air film cooling holes needing to be additionally formed in the turbine blade, and forming a new turbine blade model; and S3, calculating the new model of the turbine blade again to obtain related data, and evaluating the comprehensive cooling effect of the blade in the vicinity of the area additionally provided with the air film cooling holes. According to the method, the gas film cooling holes are additionally formed through calculation, the gas film holes are formed in the hottest position of the turbine blade to form the cooling gas film, the gas temperature of an easily-ablated area is reduced, the temperature of a local high-temperature area of the blade can be effectively reduced, thermal stress caused by high temperature is reduced, and therefore the service life of the blade is prolonged.

Description

Technical Field

[0001] The invention relates to the technical field of turbine blade cooling, in particular to a turbine blade air film cooling hole design method and a turbine blade air film cooling hole. Background Art

[0002] Film cooling is a highly efficient method for cooling turbine blades. The cold air flowing out of the film cooling holes is mixed with the mainstream combustion gas and adheres to the wall to form a cold air layer with a lower temperature. This prevents the high-temperature combustion gas from directly scouring the turbine blades and plays an important protective role for the blades. It is widely used in turbine blades of modern high-power power generation gas turbines.

[0003] Blades and end walls are two important components of gas turbine turbine cascades. Due to improper cooling design, ablation of high-temperature areas of blades often occurs. Therefore, it is necessary to add air film cooling in the prone to ablation area to reduce the gas temperature in the prone to ablation area, which can effectively reduce the temperature of the local high-temperature area of ​​the blade and reduce the thermal stress caused by high temperature, thereby increasing the service life of the blade. Summary of the invention

[0004] The purpose of this section is to summarize some aspects of embodiments of the present invention and briefly introduce some preferred embodiments. Some simplifications or omissions may be made in this section and the specification abstract and the invention title of this application to avoid blurring the purpose of this section, the specification abstract and the invention title, and such simplifications or omissions cannot be used to limit the scope of the present invention.

[0005] In view of the problem that the temperature of some areas is too high due to improper cooling design of the turbine blades, the present invention is proposed.

[0006] Therefore, an object of the present invention is to provide a method for designing film cooling holes for turbine blades.

[0007] To solve the above technical problems, the present invention provides the following technical solutions: a method for designing film cooling holes for turbine blades, comprising the following steps:

[0008] S1, modeling turbine blades with high temperature areas and obtaining relevant data;

[0009] S2, calculating the data of the film cooling holes that need to be added to the turbine blades according to the model and data obtained in S1, and forming a new turbine blade model;

[0010] S3, recalculate the new model of turbine blades to obtain relevant data and evaluate the comprehensive cooling effect of the blades in the area near the added film cooling holes.

[0011] As a preferred solution of the method for designing film cooling holes for turbine blades of the present invention, S1 includes modeling the original model of the turbine blade using modeling software to obtain a computer-readable model.

[0012] As a preferred solution of the turbine blade film cooling hole design method described in the present invention, the turbine blade original model is meshed using meshing software, and then computational fluid dynamics software is used to solve the mass equation, momentum equation, energy equation, and heat conduction equation of the original model and the temperature field of the turbine blade original model is obtained based on the above data.

[0013] As a preferred solution of the turbine blade film cooling hole design method of the present invention, wherein: S2 comprises:

[0014] S2.1 Based on the temperature field calculated in S1 and the safe operating temperature of the blade, according to the basic principle of heat transfer, using Fourier's law or other relevant heat transfer formulas, calculate the heat dissipation exceeding the safe temperature;

[0015] S2.2 According to the characteristics of the cooling gas and the principle of heat balance, the required cooling gas flow rate is calculated;

[0016] S2.3 Obtain the flow area of ​​the film cooling hole according to the cooling air flow rate and the blowing ratio.

[0017] As a preferred solution of the method for designing film cooling holes for turbine blades described in the present invention, the film cooling holes are arranged at the hottest position of the turbine blades to form a protective air film.

[0018] As a preferred solution of the turbine blade film cooling hole design method of the present invention, wherein: S2 further includes:

[0019] S2.4 obtain the number and geometric parameters of the film cooling holes according to the flow area of ​​the film cooling holes;

[0020] S2.5 uses modeling software to model the film cooling holes and obtain a new computer-readable model of the turbine blade.

[0021] As a preferred solution of the method for designing film cooling holes for turbine blades described in the present invention, the film cooling holes are selectively arranged on the blade profile, end wall, and shroud of the turbine blade, and the angles of the film cooling holes are different depending on the arrangement positions.

[0022] As a preferred solution of the turbine blade film cooling hole design method described in the present invention, wherein: in the S3, the new turbine blade model is re-gridded, and computational fluid dynamics software is used to solve the mass equation, momentum equation, energy equation, and heat conduction equation of the new model, and the temperature field of the new turbine blade model is obtained based on the above data.

[0023] As a preferred solution of the method for designing film cooling holes for turbine blades described in the present invention, it is determined whether the cooling film is effectively generated based on the temperature field of the new model of the turbine blade.

[0024] The beneficial effects of the turbine blade film cooling hole design method described in the present invention: the present invention provides the opening position of the film hole and the calculation method of the film hole structure parameters. By adding film holes in the improperly designed places of turbine blade cooling, a cooling film is formed, and the gas temperature in the easily ablated area is reduced. The temperature of the local high-temperature area of ​​the blade can be effectively reduced, and the thermal stress caused by high temperature can be reduced, thereby increasing the service life of the blade.

[0025] The present invention also provides the following technical solution: a turbine blade film cooling hole, used in the above-mentioned turbine blade film cooling hole design method, and also includes a turbine blade and a plurality of film cooling holes arranged on the turbine blade.

[0026] The beneficial effects of the turbine blade film cooling holes described in the present invention are as follows: the present invention adds film holes, transforms the film holes at the hottest part of the turbine blade to form a cooling film, thereby reducing the gas temperature in the easily ablated area, effectively reducing the temperature of the local high-temperature area of ​​the blade, reducing the thermal stress caused by high temperature, and extending the service life of the blade. BRIEF DESCRIPTION OF THE DRAWINGS

[0027] In order to more clearly illustrate the technical solutions of the embodiments of the present invention, the following briefly introduces the drawings required for describing the embodiments. Obviously, the drawings described below are only some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without creative labor. Among them:

[0028] Figure 1 This is an overall schematic diagram of a turbine blade with an internal cooling channel.

[0029] Figure 2 A cross-sectional view of a turbine blade with an internal cooling channel.

[0030] Figure 3 Temperature distribution diagram of turbine blades with internal cooling channels.

[0031] Figure 4 This is an overall schematic diagram of a turbine blade with additional film cooling holes.

[0032] Figure 5 A cross-sectional view of a turbine blade with additional film cooling holes.

[0033] Figure 6 Temperature distribution diagram of turbine blades with additional film cooling holes.

[0034] Figure 7 Schematic diagram of film cooling hole structure parameters.

[0035] Figure 8 Another schematic diagram of film cooling hole structure parameters.

[0036] Fig. 9 Schematic diagram of the belt and end wall. DETAILED DESCRIPTION

[0037] In order to make the above-mentioned objects, features and advantages of the present invention more obvious and easy to understand, the specific implementation methods of the present invention are described in detail below in conjunction with the accompanying drawings.

[0038] In the following description, many specific details are set forth to facilitate a full understanding of the present invention, but the present invention may also be implemented in other ways different from those described herein, and those skilled in the art may make similar generalizations without violating the connotation of the present invention. Therefore, the present invention is not limited to the specific embodiments disclosed below.

[0039] Secondly, the term "one embodiment" or "embodiment" as used herein refers to a specific feature, structure, or characteristic that may be included in at least one implementation of the present invention. The term "in one embodiment" that appears in different places in this specification does not necessarily refer to the same embodiment, nor does it refer to a separate or selective embodiment that is mutually exclusive with other embodiments.

[0040] Example 1

[0041] Reference Figure 1-Figure 6 According to an embodiment of the present invention, a method for designing film cooling holes for turbine blades is provided, comprising the following steps:

[0042] S1, modeling turbine blades with high temperature areas and obtaining relevant data;

[0043] By modeling the existing turbine blades (design models or drawings can be used), the original model of the turbine blade is made into a computer-readable data model, and the model is divided using meshing software. After meshing, the computational fluid dynamics software is imported to solve the mass equation, momentum equation, energy equation, and heat conduction equation of the original model, and the temperature field of the original model of the turbine blade is obtained based on the above data.

[0044] S2, calculating the data of the film cooling holes that need to be added to the turbine blades according to the model and data obtained in S1, forming a new turbine blade model, wherein the film cooling holes are arranged at the hottest position of the turbine blades to form a protective air film;

[0045] S2.1 Based on the temperature field calculated in S1 and the specified safe operating temperature of the blade, use Fourier's law or other relevant heat transfer formulas according to the basic principles of heat transfer to calculate the heat dissipation exceeding the safe temperature;

[0046] S2.2 According to the characteristics of the cooling gas (such as specific heat capacity, cooling gas temperature change, etc.), according to the principle of heat balance, that is, the heat dissipation is equal to the heat absorbed by the cooling gas, the required cooling gas flow m is calculated. c ;

[0047] S2.3 According to the cooling air flow m c and blowing ratio, and the flow area of ​​the film cooling hole is obtained;

[0048] Specifically, the flow area calculation formula of the film cooling hole in S2.2 is:

[0049]

[0050] Where: A c is the minimum flow area of ​​the film cooling hole, usually a circular hole, A c =πγ 2 , π is the circumference of a circle, γ is the radius of the smallest through-hole, generally 0.2 to 0.6 mm; m c is the cooling air flow rate outflowing from the film cooling hole, which is calculated by S2.2; M is the optimal blowing ratio, which is different at different positions of the blade and is generally 0.5 to 2; ρ ∞ is the mainstream flow density, obtained by querying the physical properties of the mainstream gas; v ∞ is the mainstream inflow velocity, obtained from the velocity field calculated by S1.

[0051] S2.4 obtain the number and geometric parameters of the film cooling holes according to the flow area of ​​the film cooling holes;

[0052] Specifically, the calculation formula for the number of film cooling holes is:

[0053]

[0054] Where: n is the number of film cooling holes; A c is the flow area of ​​the film cooling hole; π is the ratio of pi, which is 3.1415926; d is the diameter of the minimum flow hole of the film cooling hole, which is generally 0.4-1.2mm.

[0055] The geometric parameters of the film cooling hole include the diameter d of the cylindrical section of the film cooling hole, the axial length L, the angle α between the center line of the film cooling hole and the tangential plane passing through the center of the point on the blade surface, and the angle β between the center line of the film cooling hole and the spanwise plane passing through the center of the point on the blade surface.

[0056] S2.5 uses modeling software to model the film cooling holes and obtain a new computer-readable model of the turbine blade.

[0057] S3, recalculate the new model of turbine blades to obtain relevant data.

[0058] By reanalyzing the designed new model of the turbine blade, it is determined whether the modification of the film cooling hole is successful. S3 includes the following steps.

[0059] S3.1 uses modeling software to model a new model of turbine blades;

[0060] S3.2 meshes the new turbine blade model using meshing software;

[0061] S3.3 uses computational fluid dynamics software to solve the mass equation, momentum equation, energy equation, and heat conduction equation of the new turbine blade model;

[0062] 3.4 Obtain the temperature field of the new turbine blade model.

[0063] Based on the temperature field of the new turbine blade model, it is determined whether the cooling film is effectively generated.

[0064] The method can be used to calculate the position of the high-temperature area on the turbine blade, and calculate the data of the film cooling holes that need to be opened in the high-temperature area.

[0065] Example 2

[0066] refer to Figure 1-6 The present invention also provides a turbine blade film cooling hole, which is calculated using the above method and includes: a turbine blade 100 and a plurality of film cooling holes 102 arranged on the turbine blade 100.

[0067] Furthermore, the turbine blade 100 is provided with a shroud 103 at the top end and an end wall 104 at the bottom end. The turbine blade 100 is provided with a plurality of internal cooling channels 101 .

[0068] Furthermore, the film cooling hole 102 is in communication with the inner cooling channel 101 .

[0069] The film cooling holes 102 are selectively arranged on the airfoil, end wall 104, and shroud 103 of the turbine blade 100. The angles of the film cooling holes 102 are different according to the different arrangement positions, wherein the angle α is the angle between the center line of the film cooling hole and the tangential plane passing through the center of the point on the blade surface, the angle β is the angle between the center line of the film cooling hole and the spanwise plane passing through the center of the point on the blade surface, d is the diameter of the cylindrical section of the film cooling hole, and L is the axial length of the film cooling hole.

[0070] The film cooling holes 102 arranged on the back side of the blade have an α angle between 25° and 35°, and a β angle between 60° and 90°.

[0071] The film cooling holes 102 arranged on the blade basin side have an α angle between 25° and 45°, and a β angle between 30° and 90°.

[0072] The film cooling holes 102 arranged on the end wall 104 have an α angle between 25° and 45°, and a β angle between 30° and 90°.

[0073] Reference Figure 1-Figure 3 , which is a model without two cooling holes on the blade body, the average temperature of the local area of ​​the turbine blade is 844℃, and the maximum temperature is 858℃.

[0074] Reference Figure 4-Figure 6 , which is a model with two cooling holes added to the blade body. The local area where the cooling holes are located (with Figure 3 The high temperature area is located in the same position) with an average temperature of 815℃ and a maximum temperature of 842℃.

[0075] In the prior art, there is a cooling channel inside the blade, and there is an air film cooling hole on the leading edge of the blade. The high-temperature mainstream gas flows through the turbine moving blade, and the leading edge and the blade basin side are close to the leading edge. The blade is well cooled, but there is a high-temperature area near the trailing edge on the blade basin side at the top of the blade ( Figure 3 In this embodiment, two cooling holes are added near the trailing edge of the blade basin at the top of the blade, and the average temperature of the high temperature area is reduced by 29°C, and the maximum temperature is reduced by 16°C, which effectively reduces the temperature of the high temperature area of ​​the turbine blade.

[0076] Importantly, it should be noted that the construction and arrangement of the present application shown in a plurality of different exemplary embodiments are only exemplary. Although only a few embodiments are described in detail in this disclosure, it should be readily understood by those who refer to this disclosure that many modifications are possible (e.g., the size, scale, structure, shape and proportion of various elements, and parameter values ​​(e.g., temperature, pressure, etc.), installation arrangement, use of materials, color, directional changes, etc.) without substantially departing from the novel teachings and advantages of the subject matter described in the application. For example, the element shown as integrally formed can be composed of multiple parts or elements, the position of the element can be inverted or otherwise changed, and the nature or number or position of the discrete element can be changed or changed. Therefore, all such modifications are intended to be included in the scope of the present invention. The order or sequence of any process or method steps can be changed or reordered according to alternative embodiments. In the claims, any "device plus function" clause is intended to cover the structure of performing the function described herein, and is not only structurally equivalent but also equivalent structure. Without departing from the scope of the present invention, other replacements, modifications, changes and omissions can be made in the design, operating conditions and arrangement of the exemplary embodiments. Therefore, the invention is not limited to a specific embodiment, but extends to numerous modifications still falling within the scope of the appended claims.

[0077] Additionally, in order to provide a concise description of exemplary embodiments, all features of an actual embodiment (ie, those features that are not relevant to the best mode presently contemplated for carrying out the invention or those that are not relevant to implementing the invention) may not be described.

[0078] It will be appreciated that in the development of any actual implementation, as in any engineering or design project, numerous implementation-specific decisions may be made. Such a development effort may be complex and time-consuming, but will be a routine task of design, fabrication, and production for those of ordinary skill having the benefit of this disclosure without undue experimentation.

[0079] It should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention rather than to limit it. Although the present invention has been described in detail with reference to the preferred embodiments, those skilled in the art should understand that the technical solutions of the present invention may be modified or replaced by equivalents without departing from the spirit and scope of the technical solutions of the present invention, which should all be included in the scope of the claims of the present invention.

Claims

1. A method for designing film cooling holes for turbine blades, characterized in that: The following steps are included: S1, modeling turbine blades with high temperature areas and obtaining relevant data; S2, calculating the data of the film cooling holes that need to be added to the turbine blades according to the model and data obtained in S1, and forming a new turbine blade model; S3, recalculate the new model of turbine blades to obtain relevant data and evaluate the comprehensive cooling effect of the blades in the area near the added film cooling holes.

2. The method for designing film cooling holes for turbine blades according to claim 1, characterized in that: The S1 includes modeling the original model of the turbine blade using modeling software to obtain a computer-readable model.

3. The method for designing film cooling holes for turbine blades according to claim 2, characterized in that: The grid division software is used to divide the original model of the turbine blade into grids, and then the computational fluid dynamics software is used to solve the mass equation, momentum equation, energy equation, and heat conduction equation of the original model, and the temperature field of the original model of the turbine blade is obtained based on the above data.

4. The method for designing film cooling holes for turbine blades according to claim 3, characterized in that: The S2 comprises, S2.1 Based on the temperature field calculated in S1 and the safe operating temperature of the blade, according to the basic principle of heat transfer, using Fourier's law or other relevant heat transfer formulas, calculate the heat dissipation exceeding the safe temperature; S2.2 According to the characteristics of the cooling gas and the principle of heat balance, the required cooling gas flow rate is calculated; S2.3 Obtain the flow area of ​​the film cooling hole according to the cooling air flow rate and the blowing ratio.

5. The method for designing film cooling holes for turbine blades according to claim 4, characterized in that: The air film cooling holes are arranged at the hottest position of the turbine blades to form a protective air film.

6. The method for designing film cooling holes for turbine blades according to claim 5, characterized in that: Said S2 further comprises: S2.4 obtain the number and geometric parameters of the film cooling holes according to the flow area of ​​the film cooling holes; S2.5 uses modeling software to model the film cooling holes and obtain a new computer-readable model of the turbine blade.

7. The method for designing film cooling holes for turbine blades according to claim 6, characterized in that: The film cooling holes are selectively arranged on the blade profile, end wall and shroud of the turbine blade, and the angles of the film cooling holes are different according to different arrangement positions.

8. The method for designing film cooling holes for turbine blades according to claim 7, characterized in that: In the S3, the new turbine blade model is re-gridified, and computational fluid dynamics software is used to solve the mass equation, momentum equation, energy equation, and heat conduction equation of the new model, and the temperature field of the new turbine blade model is obtained based on the above data.

9. The method for designing film cooling holes for turbine blades according to claim 8, characterized in that: Based on the temperature field of the new turbine blade model, it is determined whether the cooling film is effectively generated.

10. A turbine blade film cooling hole, calculated using the turbine blade film cooling hole design method according to any one of claims 1 to 9, characterized in that: The invention comprises a turbine blade (100) and a plurality of film cooling holes (102) arranged on the turbine blade (100).

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