A turbine blade for a gas turbine

Abstract

The utility model discloses a turbine blade of gas turbine, including the cooling cavity chamber of multiple interval distribution from below to above inside turbine blade and multiple cooling channel that penetrates turbine blade from bottom to top for the cooling inside turbine blade of gas turbine, based on the design of turbine blade cooling channel and cooling cavity chamber of the utility model, cooling gas can enter from the port of turbine blade bottom cooling channel, from below to above fills different height cooling cavity in proper order, and flows out through the port of turbine blade top cooling channel, compared with the blade of traditional cooling channel, the interior heat exchange area becomes big, improves the cooling effect and gas turbine overall efficiency, can effectively promote the cooling performance, and can also give consideration to manufacturing feasibility and economy.

Description

Technical Field

[0001] This utility model relates to the field of gas turbine cooling technology, and in particular to a turbine blade for a gas turbine. Background Technology

[0002] As a core component of modern aero engines and heavy industrial applications, gas turbines play a crucial role in energy conversion efficiency, power output, and environmental protection. Among these components, turbine blades are key parts of the gas turbine that withstand the highest temperatures and stresses. Operating in extreme high-temperature environments, turbine blades face severe thermal load challenges, placing extremely high demands on the selection of blade materials and cooling design.

[0003] Traditional turbine blade cooling primarily relies on the design of its internal cooling channels to reduce the blade's operating temperature. These cooling channels are typically circular in cross-section, and cooling is achieved by introducing cooling air from the compressor that flows through the interior of the blade. While this design could meet the needs of early gas turbines to some extent, its limitations have become increasingly apparent with the continuous improvement of engine efficiency, power density, and environmental protection requirements.

[0004] Due to the limited space inside the blades, traditional circular cross-section cooling channels cannot provide sufficient surface area for heat exchange, limiting further improvements in cooling efficiency. To ensure adequate cooling, it is often necessary to increase the cooling airflow, but this leads to a decrease in the overall efficiency of the gas turbine.

[0005] In recent years, the industry has attempted to improve turbine blades using irregularly shaped channels and turbulence structures, aiming to enhance cooling performance while reducing the impact on overall engine efficiency. However, the results have often been unsatisfactory, limiting their widespread application in practical engineering. Utility Model Content

[0006] The purpose of this invention is to overcome the shortcomings of the prior art and provide a turbine blade for a gas turbine that, in addition to having a cooling channel, also has multiple cooling chambers spaced apart from bottom to top, which can improve cooling performance while reducing the impact on the overall efficiency of the engine.

[0007] To achieve the above objectives, this utility model employs the following technical solution:

[0008] This utility model provides a turbine blade for a gas turbine, including multiple cooling chambers spaced apart from bottom to top inside the turbine blade, and multiple cooling channels penetrating the turbine blade from bottom to top, for cooling the interior of the gas turbine blade; the cross-section of the cooling channels is circular or elliptical; the inner wall of the cooling channels is also provided with rib structures; multiple frustums are provided inside the cooling chambers; and multiple impact cooling holes are opened on the outer wall of the turbine blade corresponding to the location of the cooling chambers.

[0009] Optionally, the cooling channels are arranged along the mid-arc line of the turbine blade profile.

[0010] Optionally, the ratio of the major axis to the minor axis of the ellipse is between 1.2:1 and 1.5:1.

[0011] Optionally, the taper of the frustum is 2°-4°.

[0012] Optionally, the plurality of impact cooling holes are evenly distributed along the circumference of the cooling chamber.

[0013] Optionally, the diameters of the plurality of impact cooling holes are configured with differentiated spacing.

[0014] Optionally, the diameter of the impact cooling hole is 1 mm to 2 mm.

[0015] Optionally, the thickness of the outer wall of the turbine blade corresponding to the location of the cooling chamber is 3%-6% of the maximum thickness of the turbine blade profile.

[0016] Optionally, the height of the cooling chamber is 2%-6% of the gas turbine blade height.

[0017] Compared with the prior art, the beneficial effects achieved by this utility model are as follows:

[0018] This invention proposes a turbine blade for a gas turbine. The turbine blade has multiple cooling channels running from bottom to top and multiple cooling chambers spaced apart from bottom to top. Based on the design of the cooling channels and cooling chambers of this invention, cooling gas can enter from the port of the cooling channel at the bottom of the turbine blade, fill the cooling chambers at different heights from bottom to top, and then flow out through the port of the cooling channel at the top of the turbine blade. Compared with blades with traditional cooling channels, the internal heat exchange area is increased, which improves the cooling effect and the overall efficiency of the gas turbine. It can effectively improve cooling performance while taking into account manufacturing feasibility and economy.

[0019] This utility model proposes a turbine blade for a gas turbine, wherein the cooling channel can be a circular cross-section or an elliptical cross-section. The elliptical cross-section cooling channel can increase the flow rate of the cooling medium per unit area and optimize the distribution of the cooling medium inside the blade. The ratio of the major and minor axes of the ellipse can be precisely controlled by the operator to further adjust the cooling effect and meet the cooling requirements under different operating conditions. In addition, a fin structure is introduced to further increase the heat exchange area and enhance the cooling effect.

[0020] This invention proposes a turbine blade for a gas turbine. The design of the turbine blade cooling channel and cooling chamber can effectively improve the heat dissipation performance of the turbine blade without significantly increasing the blade weight, thereby improving the overall efficiency and reliability of the gas turbine.

[0021] The present invention proposes a turbine blade for a gas turbine, wherein the outer wall of the blade is provided with impact cooling holes connected to the cooling chamber. The impact cooling holes can cool the outer wall of the blade, effectively increasing the heat dissipation performance and cooling efficiency of the turbine blade. The diameter of the impact cooling holes is configured with differentiated intervals, specifically, the adjacent holes exhibit a periodic alternating distribution of large and small diameters, thereby optimizing the flow and heat exchange effect of the cooling medium.

[0022] The present invention proposes a turbine blade for a gas turbine, wherein multiple frustums arranged in the shape of turbine blades are provided in the cooling chamber, which can ensure the strength of the blades and rationally distribute the cooling flow, thereby increasing the cooling effect of the turbine blades. Attached Figure Description

[0023] To more clearly illustrate the technical solutions in the embodiments of this utility model or the prior art, the drawings used in the embodiments will be briefly described below. Obviously, the drawings described below are only some embodiments of this utility model. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort, wherein:

[0024] Figure 1 The image shown is a front view of the turbine blades of a gas turbine in one embodiment of this utility model;

[0025] Figure 2 The image shown is a top view of the turbine blades of a gas turbine in one embodiment of this utility model;

[0026] Figure 3 As shown Figure 1 AA section view;

[0027] Figure 4 As shown Figure 1 BB section view;

[0028] Figure 5 As shown Figure 4 A magnified view of part C.

[0029] In the diagram: 1. Turbine blade; 2. Impact cooling hole; 3. Cooling channel with elliptical cross section; 4. Rib structure; 5. Frustum; 6. Cooling chamber. Detailed Implementation

[0030] The present invention will be further described below with reference to the accompanying drawings. The following embodiments are only used to more clearly illustrate the technical solution of the present invention, and should not be used to limit the scope of protection of the present invention.

[0031] In the description of this utility model, it should be understood that the terms "center," "longitudinal," "lateral," "upper," "lower," "front," "rear," "left," "right," "vertical," "horizontal," "top," "bottom," "inner," and "outer," etc., indicating orientation or positional relationships, are based on the orientation or positional relationships shown in the accompanying drawings and are only for the convenience of describing this utility model and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation, and therefore should not be construed as a limitation of this utility model. Furthermore, the terms "first," "second," etc., are used for descriptive purposes only and should not be construed as indicating or implying relative importance or implicitly specifying the number of indicated technical features. Thus, features defined with "first," "second," etc., may explicitly or implicitly include one or more of that feature. In the description of this utility model, unless otherwise stated, "a plurality of" means two or more.

[0032] In the description of this utility model, it should be noted that, unless otherwise explicitly specified and limited, the terms "installation," "connection," and "joining" should be interpreted broadly. For example, they can refer to a fixed connection, a detachable connection, or an integral connection; they can refer to a mechanical connection or an electrical connection; they can refer to a direct connection or an indirect connection through an intermediate medium; and they can refer to the internal connection of two components. Those skilled in the art can understand the specific meaning of the above terms in this utility model based on the specific circumstances.

[0033] The application principle of this utility model will be described in detail below with reference to the accompanying drawings.

[0034] like Figure 1-5 As shown in the figure, the present invention provides a turbine blade for a gas turbine, including multiple cooling chambers 6 distributed from bottom to top (from bottom to top) inside the turbine blade 1, and multiple cooling channels 3 with elliptical cross sections running through the turbine blade from bottom to top, for guiding the flow of cooling medium to reduce the operating temperature of the turbine blade 1.

[0035] In this embodiment, the elliptical cross-section cooling channels 3 are arranged along the mid-arc line of the turbine blade 1, such as... Figure 2 and Figure 3 As shown, eight cooling channels 3 with elliptical cross sections are arranged along the mid-arc line of the turbine blade 1, and the direction of the cooling channels 3 is based on the blade height direction of the turbine blade 1; the cross-sectional size of the cooling channels 3 with elliptical cross sections is adjusted appropriately according to the size of the turbine blade.

[0036] In this embodiment, the cooling channel 3 with an elliptical cross-section has a major axis to minor axis ratio between 1.2:1 and 1.5:1 to optimize the flow rate and flow distribution of the cooling medium.

[0037] In this embodiment, the cooling channel 3 with an elliptical cross-section is also provided with a rib structure 4. The rib structure is set at the major axis end of the cooling channel 3 with an elliptical cross-section (the major axis end of the elliptical cross-section), which aims to maximize the heat exchange area and improve the cooling efficiency, while taking into account the flow resistance and structural reliability. Specifically, this embodiment does not limit the shape of the rib structure 4. Straight ribs, annular ribs, trapezoidal ribs, etc. can all be provided.

[0038] In this embodiment, the cooling chamber 6 can be regarded as a chamber formed by two upper and lower leaf-shaped intermediate walls and a ring of outer walls;

[0039] For details, please refer to Figure 4 The number of cooling chambers 6 is set to nine, namely the first, second, third, fourth, fifth, sixth, seventh, eighth and ninth cooling chambers. The elliptical cross-section cooling channel 3 passes through the first intermediate wall, the first cooling chamber, the second intermediate wall, the second cooling chamber, the third intermediate wall, the third cooling chamber, the fourth intermediate wall, the fourth cooling chamber, the fifth intermediate wall, the fifth cooling chamber, the sixth intermediate wall, the sixth cooling chamber, the seventh intermediate wall, the seventh cooling chamber, the eighth intermediate wall, the eighth cooling chamber, the ninth intermediate wall, the ninth cooling chamber and the tenth intermediate wall from bottom to top. The lower surface of the first intermediate wall is the bottom surface of the turbine blade 1, and the upper surface of the tenth intermediate wall is the top surface of the turbine blade 1.

[0040] Specifically, the height of cooling chamber 6 is 2%-6% of the height of the gas turbine blades.

[0041] In the embodiment, reference Figure 3 and Figure 5 The cooling chamber 6 is provided with multiple frustums 5 to ensure the overall strength of the turbine blade 1 and enhance the cooling effect on the blade; the upper and lower surfaces of the frustums 5 are connected to two adjacent intermediate walls of the cooling chamber 6, and the taper of the frustums 5 is 2°-4°.

[0042] In this embodiment, reference Figure 1 , Figure 3 , Figure 4 , Figure 5 Multiple impact cooling holes 2 are formed on the outer wall of the turbine blade corresponding to the location of the cooling chamber 6, and these multiple impact cooling holes 2 are evenly distributed along the circumference of the cooling chamber 6. Specifically, a ring of impact cooling holes 2 is set at the center line of the outer wall surrounding the cooling chamber 6. The diameter of the impact cooling holes 2 is set to 1 mm-2 mm and is manually set. The diameter of the impact cooling holes 2 is configured with differentiated intervals. Specifically, adjacent holes exhibit a periodic alternation of large and small diameters to optimize the flow and heat exchange effect of the cooling medium. Specifically, the thickness of the outer wall of the turbine blade 1 corresponding to the location of the cooling chamber 6 is 3%-6% of the maximum thickness of the turbine blade 1.

[0043] In this embodiment, based on the structure of a gas turbine blade described above, the working process of the turbine blade in this embodiment is provided as follows:

[0044] First, cooling gas is extracted from the compressor, output through the gas turbine shaft system to the bottom of the gas turbine blades, and enters from the port of the cooling channel at the bottom of the turbine blades;

[0045] Secondly, after the cooling gas enters the cooling channel, it fills the cooling chambers at different heights from bottom to top. In the cooling chamber inside the turbine blade, there are multiple truncated cones that increase disturbance. When the cooling gas flows through the truncated cones, it will generate vortices and turbulence, further breaking the laminar flow state and allowing the cooling gas to fully contact and exchange heat with the inner wall of the turbine blade.

[0046] Subsequently, the cooling gas in the cooling chamber is injected vertically at high speed onto the outer surface of the blade through a ring of impact cooling holes on the outer wall, forming a local strong cooling zone at the impact point, effectively removing heat and cooling the surface of the turbine blade.

[0047] Finally, about one-third of the cooling gas flows out from the port of the cooling channel at the top of the turbine blade, forming a low-temperature protective gas film at the top edge of the turbine blade, which fully cools the top of the turbine blade.

[0048] In summary, based on a turbine blade structure for a gas turbine, its combined cooling method of "internal turbulence + surface impact + top film cooling" can improve cooling performance while reducing the impact on the overall engine efficiency, and can ensure that the turbine blade can still work reliably in high-temperature environments above 1200℃.

[0049] The embodiments of the present utility model have been described above with reference to the accompanying drawings. However, the present utility model is not limited to the specific embodiments described above. The specific embodiments described above are merely illustrative and not restrictive. Those skilled in the art can make many other forms under the guidance of the present utility model without departing from the spirit and scope of the claims. All of these forms are within the protection scope of the present utility model.

[0050] The foregoing has shown and described the basic principles, main features, and advantages of this utility model. Those skilled in the art should understand that this utility model is not limited to the above embodiments. The embodiments and descriptions in the specification are merely illustrative of the principles of this utility model. Various changes and modifications can be made to this utility model without departing from its spirit and scope, and all such changes and modifications fall within the scope of the claims. The scope of protection of this utility model is defined by the appended claims and their equivalents.

Claims

1. A turbine blade for a gas turbine, characterized in that, The device includes multiple cooling chambers spaced from bottom to top inside the turbine blade, and multiple cooling channels running from bottom to top through the turbine blade, used to cool the inside of the gas turbine blade; the cross-section of the cooling channels is circular or elliptical; the inner wall of the cooling channels is also provided with rib structures; multiple frustums are provided inside the cooling chambers; and multiple impact cooling holes are opened on the outer wall of the turbine blade corresponding to the location of the cooling chambers.

2. The turbine blade of the gas turbine according to claim 1, characterized in that, The cooling channels are distributed and arranged along the mid-arc line of the turbine blade profile.

3. The turbine blade of the gas turbine according to claim 1, characterized in that, The ratio of the major axis to the minor axis of the ellipse is between 1.2:1 and 1.5:

1.

4. The turbine blade of the gas turbine according to claim 1, characterized in that, The taper of the frustum is 2°-4°.

5. The turbine blade of the gas turbine according to claim 1, characterized in that, The plurality of impact cooling holes are evenly distributed along the perimeter of the cooling chamber.

6. The turbine blade of the gas turbine according to claim 1, characterized in that, The diameters of the plurality of impact cooling holes are configured with differentiated intervals.

7. The turbine blade of the gas turbine according to claim 6, characterized in that, The diameter of the impact cooling hole is 1 mm-2 mm.

8. The turbine blade of the gas turbine according to claim 1, characterized in that, The thickness of the outer wall of the turbine blade corresponding to the location of the cooling chamber is 3%-6% of the maximum thickness of the turbine blade profile.

9. The turbine blade of the gas turbine according to claim 1, characterized in that, The height of the cooling chamber is 2%-6% of the height of the gas turbine blades.

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