Cooling structure for front edge of turbine blade

Abstract

The invention provides a cooling structure for the front edge of a turbine blade, the cooling structure comprises an air film hole structure and an impact hole structure, the air film hole structure is arranged on the wall surface of the end part of the front edge of the turbine blade, and the impact hole structure is arranged on a front edge impact plate of the turbine blade and is used for communicating a cold air cavity with an impact cavity. The impact hole structure comprises a plurality of first oblique impact holes and second oblique impact holes which are arranged in a crossed mode in the blade height direction, and impact stagnation points of the first oblique impact holes and the second oblique impact holes are located on the inner wall faces of the front edges of the turbine blades on the two sides of the air film hole structure correspondingly. According to the cooling structure, the convective heat transfer coefficient of the inner wall face of the front edge of the turbine blade can be increased, the target surface area influenced by impact is increased, the temperature uniformity of the wall face of the front edge is improved, the cooling effect near the front edge of the turbine blade is greatly improved, and the using amount of cooling gas is reduced; and a positive help effect is achieved on improving the performance and the efficiency of the engine.

Description

Technical Field

[0001] The invention belongs to the field of turbines, which are high-temperature components of gas turbines, and relates to a cooling design technology for the leading edge of turbine blades, in particular to a cooling structure for the leading edge of turbine blades. Background Art

[0002] Gas turbine engines are thermal power devices with powerful output power and high thermal efficiency. They are usually in extremely harsh working environments and have three characteristics: high temperature, high pressure and high speed. They are now widely used in modern military and industry. In the process of developing gas turbine engines, in order to improve the performance of gas turbine engines, the temperature before the turbine is continuously increased. For example, the temperature of the turbine inlet gas of the current advanced aviation turbofan engine has reached 1800K~2050K, which greatly exceeds the temperature bearing capacity of existing high-temperature alloys. The leading edge of the turbine blades, which is directly impacted by the high-temperature gas, faces greater challenges due to the higher temperature it withstands. In order to ensure the safe operation of the turbine, effective cooling must be implemented to reduce the metal wall temperature of the leading edge of the high-temperature turbine blades.

[0003] At present, cooling structures such as impingement cooling, air film cooling, and impingement air film combined cooling are usually used to cool the leading edge of turbine blades to protect the long-term safe, reliable and efficient operation of gas turbine engines. Existing impingement holes include direct impingement, swirl impingement, cylindrical impingement holes, and conical impingement holes. Both direct impingement and swirl impingement have the problem of uneven cooling of the inner wall surface. The uneven temperature distribution means a large temperature gradient, which can easily cause thermal fatigue fracture of the blade. Although swirl impingement can improve the uniformity of the temperature distribution on the blade wall to a certain extent, the average cooling efficiency of the entire wall surface will decrease, and good cooling performance cannot be achieved.

[0004] Therefore, there is an urgent need for a new structure that can achieve efficient cooling of the leading edge of the turbine blade while improving the uniformity of the wall temperature distribution. Summary of the invention

[0005] In order to solve the technical problems of uneven cooling of existing impact holes leading to thermal fatigue fracture of blades and low cooling performance, the present invention discloses a cooling structure for the leading edge of a turbine blade, the cooling structure comprising an air film hole structure and an impact hole structure, the air film hole structure is arranged on the wall surface of the leading edge end of the turbine blade, and the impact hole structure is arranged on the leading edge impact plate of the turbine blade, for connecting the cold air cavity and the impact cavity.

[0006] Among them, the impact hole structure includes a plurality of first oblique impact holes and second oblique impact holes arranged crosswise along the blade height direction, and the impact stagnation points of the first oblique impact holes and the second oblique impact holes are respectively located on the suction side and the pressure side of the inner wall surface of the leading edge of the turbine blade on both sides of the air film hole structure.

[0007] Furthermore, the angles between the first oblique impact hole and the second oblique impact hole and the normal line of the leading edge impact plate are in the range of 30° to 45°.

[0008] Furthermore, the first oblique impact hole and the second oblique impact hole have the same included angle with the normal line of the leading edge impact plate.

[0009] Furthermore, the first oblique impact hole and the second oblique impact hole are hourglass-shaped structures that are wide at both ends and narrow in the middle.

[0010] Furthermore, the hourglass-shaped structure includes a cylindrical section, and both ends of the cylindrical section are respectively provided with conical sections with gradually increasing diameters.

[0011] Preferably, the maximum diameter of the conical cylinder section is 1.2 to 2.0 times the diameter of the cylindrical section, wherein the maximum diameter is the diameter of the conical cylinder section at the position with the shortest length along the normal direction of the leading edge impact plate.

[0012] Preferably, the axis of the cylindrical section coincides with the axis of the conical barrel section.

[0013] Preferably, the length ratio of the cylindrical section to the conical section is 0.5-1.5.

[0014] Preferably, the length of the cylindrical segment is 2 to 4 times the diameter of the cylindrical segment.

[0015] Furthermore, the film hole structure is a multi-row and multi-column film hole structure, and the number of leading edge film holes in each row arranged along the blade height direction is equal to the sum of the number of the first oblique impact holes and the second oblique impact holes.

[0016] Compared with the prior art, the beneficial effects that can be achieved by at least one of the above-mentioned technical solutions adopted in the embodiments of this specification include at least the following: in the present invention, the impact hole structure is designed as a plurality of first oblique impact holes and second oblique impact holes arranged crosswise along the blade height direction, which can enhance the convective heat transfer coefficient of the inner wall surface of the leading edge of the turbine blade, increase the target surface area affected by the impact, improve the temperature uniformity of the leading edge wall surface, greatly improve the cooling effect near the leading edge of the turbine blade, reduce the amount of cooling gas used, and play a positive role in improving the performance and efficiency of the engine. BRIEF DESCRIPTION OF THE DRAWINGS

[0017] In order to more clearly illustrate the technical solutions of the embodiments of the present application, the drawings required for use in the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present application. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying creative work.

[0018] Figure 1 A three-dimensional diagram of a cooling structure on the leading edge of a turbine blade disclosed in an embodiment of the present invention;

[0019] Figure 2 A cross-sectional view of a cooling structure on the leading edge of a turbine blade disclosed in an embodiment of the present invention;

[0020] Figure 3 A schematic diagram of an hourglass-shaped structure with wide ends and narrow middle disclosed in an embodiment of the present invention;

[0021] Among them, 1. cold air cavity; 2. leading edge impact plate; 3. first oblique impact hole; 4. second oblique impact hole; 5. impact cavity; 6. leading edge end of turbine blade; 7. leading edge air film hole; 8. second impact stagnation point; 9. first impact stagnation point; 10. cylindrical section; 20. conical tube section. DETAILED DESCRIPTION

[0022] The embodiments of the present application are described in detail below with reference to the accompanying drawings.

[0023] The following describes the implementation methods of the present application through specific specific examples, and those skilled in the art can easily understand other advantages and effects of the present application from the content disclosed in this specification. Obviously, the described embodiments are only a part of the embodiments of the present application, rather than all the embodiments. The present application can also be implemented or applied through other different specific implementation methods, and the details in this specification can also be modified or changed in various ways based on different viewpoints and applications without departing from the spirit of the present application. It should be noted that the following embodiments and the features of the embodiments can be combined with each other in the absence of conflict. Based on the embodiments in the present application, all other embodiments obtained by ordinary technicians in the field without making creative work belong to the scope of protection of the present application.

[0024] In the description of this embodiment, it should be understood that the terms "center", "longitudinal", "lateral", "up", "down", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inside" and "outside" etc., indicating orientations or positional relationships, are based on the orientations or positional relationships shown in the accompanying drawings, and are only for the convenience of describing the invention and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and therefore should not be understood as a limitation on the invention.

[0025] In addition, the terms "first", "second", "third", etc. are used for descriptive purposes only and should not be understood as indicating or implying relative importance or implicitly indicating the number of technical features indicated. Therefore, features defined as "first", "second", etc. may explicitly or implicitly include one or more features. In the description of the present invention, unless otherwise specified, the meaning of "plurality" is two or more.

[0026] It should also be noted that the illustrations provided in the following embodiments are only schematic illustrations of the basic concept of the present disclosure. The drawings only show components related to the present disclosure rather than being drawn according to the number, shape and size of components in actual implementation. In actual implementation, the type, quantity and proportion of each component may be changed arbitrarily, and the component layout may also be more complicated.

[0027] The present invention discloses a cooling structure for the leading edge of a turbine blade, the cooling structure comprising an air film hole structure and an impact hole structure, see Figure 1 As shown, the film hole structure is arranged on the wall surface of the leading edge end 6 of the turbine blade, and the film hole structure is a multi-row and multi-column film hole structure, and each row and each column includes a plurality of leading edge film holes 7. The impact hole structure is arranged on the leading edge impact plate 2 of the turbine blade, and is used to connect the cold air cavity 1 and the impact cavity 5.

[0028] Among them, see Figure 1 and Figure 2 As shown, the impact hole structure includes a plurality of first oblique impact holes 3 and second oblique impact holes 4 arranged crosswise along the blade height direction, and the impact stagnation points of the first oblique impact holes 3 and the second oblique impact holes 4 are respectively located on the suction side and the pressure side of the inner wall surface of the leading edge of the turbine blade on both sides of the air film hole structure.

[0029] When the gas turbine engine is working, the secondary flow gas enters the cold air cavity 1 of the turbine blade from the blade root, enters the first oblique impact hole 3 and the second oblique impact hole 4, and is accelerated and then output to the impact cavity 5 respectively, and impacts the suction side and the leading edge pressure side of the leading edge wall of the turbine blade respectively. During the development process, the wall jet conducts convection heat exchange with the leading edge inner wall, and then flows out from the leading edge film hole 7 when reaching the leading edge film hole 7 of the film hole structure, thereby protecting the leading edge outer wall. The cross-arranged first oblique impact hole 3 and second oblique impact hole 4 in the present invention can increase the speed of the secondary flow gas, increase the momentum of the impact target surface, and thus enhance the forced convection heat exchange effect between the cold air and the inner wall. At the same time, by arranging the first oblique impact hole 3 and the second oblique impact hole 4, the ejected secondary flow gas has a certain angle with the normal of the leading edge impact plate 2, so that the impact stagnation point deviates from the stagnation line of the turbine blade, and the wall jet can be more fully developed, thereby affecting a larger target surface area, and the area of ​​high convective heat transfer coefficient is increased, thereby improving the temperature uniformity of the leading edge wall. The heat exchange effect of the inner wall affects the entire leading edge wall along the wall thickness direction after heat conduction through the leading edge wall of the turbine blade, thereby achieving the effect of reducing the leading edge wall temperature of the turbine blade, and making the leading edge wall temperature distribution more uniform, thereby reducing thermal stress.

[0030] Further, the angle between the first oblique impact hole 3 and the second oblique impact hole 4 and the normal of the leading edge impact plate 2 is in the range of 30° to 45°, so as to ensure that the impact stagnation points of the first oblique impact hole 3 and the second oblique impact hole 4 are on both sides of the air film hole structure. Figure 2 As shown, the angle between the second oblique impact hole 4 and the normal of the leading edge impact plate 2 is β, and its corresponding impact stagnation point is the first impact stagnation point 9. The angle between the first oblique impact hole 3 and the normal of the leading edge impact plate 2 is α, and its corresponding impact stagnation point is the second impact stagnation point 8.

[0031] Furthermore, the first oblique impact hole 3 and the second oblique impact hole 4 have the same included angle with the normal line of the leading edge impact plate 2 , that is, α=β.

[0032] Furthermore, the first oblique impact hole 3 and the second oblique impact hole 4 are hourglass-shaped structures that are wide at both ends and narrow in the middle.

[0033] Further, see Figure 3 As shown, the hourglass-like structure includes a cylindrical section 10, and both ends of the cylindrical section 10 are respectively provided with a conical section 20 with a gradually increasing diameter. The transition between the cylindrical section 10 and the conical section 20 should be smooth, and no steps should be formed in the hole.

[0034] Preferably, see Figure 3As shown, the maximum diameter of the cone section 20 is 1.2 to 2.0 times the diameter d1 of the cylindrical section 10, wherein the maximum diameter is the diameter of the cone section 20 at the minimum length position along the normal direction of the leading edge impact plate, that is, Figure 3 At d2 shown in the figure, the cold air is accelerated in the impact hole, but the resistance loss caused by flow separation due to sudden contraction or expansion is avoided.

[0035] Preferably, the axis of the cylindrical section 10 coincides with the axis of the conical barrel section 20 .

[0036] Preferably, the length ratio of the cylindrical section 10 to the conical section 20 is 0.5-1.5.

[0037] Preferably, the length of the cylindrical segment 10 is 2 to 4 times the diameter of the cylindrical segment 10 .

[0038] Furthermore, the film hole structure is a multi-row and multi-column film hole structure, and the number of leading edge film holes 7 in each row arranged along the blade height direction is equal to the sum of the number of the first oblique impact holes 3 and the second oblique impact holes 4 .

[0039] The embodiments of the present invention achieve the following technical effects: in the present invention, the impact hole structure is designed as a plurality of first oblique impact holes and second oblique impact holes arranged crosswise along the blade height direction, which can enhance the convective heat transfer coefficient of the inner wall surface of the leading edge of the turbine blade, increase the target area affected by the impact, improve the temperature uniformity of the leading edge wall surface, greatly improve the cooling effect near the leading edge of the turbine blade, reduce the amount of cooling gas used, and play a positive role in improving the performance and efficiency of the engine.

[0040] Obviously, those skilled in the art should understand that the above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. For those skilled in the art, the embodiments of the present invention may have various modifications and changes. Any modification, equivalent replacement, improvement, etc. made within the spirit and principle of the present invention shall be included in the protection scope of the present invention.

[0041] In addition, it should be understood that although the present specification is described according to implementation modes, not every implementation mode contains only one independent technical solution. This description of the specification is only for the sake of clarity. Those skilled in the art should regard the specification as a whole. The technical solutions in each embodiment may also be appropriately combined to form other implementation modes that can be understood by those skilled in the art.

Claims

1. A cooling structure for the leading edge of a turbine blade, characterized in that: It comprises an air film hole structure and an impact hole structure, wherein the air film hole structure is arranged on the wall surface of the leading edge end (6) of the turbine blade, and the impact hole structure is arranged on the leading edge impact plate (2) of the turbine blade, and is used to connect the cold air cavity (1) and the impact cavity (5); The impact hole structure comprises a plurality of first oblique impact holes (3) and second oblique impact holes (4) arranged crosswise along the blade height direction, and the impact stagnation points of the first oblique impact holes (3) and the second oblique impact holes (4) are respectively located on the suction side and the pressure side of the inner wall surface of the leading edge of the turbine blade on both sides of the air film hole structure.

2. The cooling structure for the leading edge of a turbine blade according to claim 1, characterized in that: The included angle between the first oblique impact hole (3) and the second oblique impact hole (4) and the normal line of the leading edge impact plate (2) is in the range of 30° to 45°.

3. The cooling structure for the leading edge of a turbine blade according to claim 2, characterized in that: The first oblique impact hole (3) and the second oblique impact hole (4) have the same included angle with the normal line of the leading edge impact plate (2).

4. The cooling structure for the leading edge of a turbine blade according to claim 1, characterized in that: The first oblique impact hole (3) and the second oblique impact hole (4) are hourglass-shaped structures that are wide at both ends and narrow in the middle.

5. The cooling structure for the leading edge of a turbine blade according to claim 4, characterized in that: The hourglass-shaped structure comprises a cylindrical section (10), and both ends of the cylindrical section (10) are respectively provided with conical cylinder sections (20) with gradually increasing diameters.

6. The cooling structure for the leading edge of a turbine blade according to claim 5, characterized in that: The maximum diameter of the conical cylinder section (20) is 1.2 to 2.0 times the diameter of the cylindrical section (10), wherein the maximum diameter is the diameter of the conical cylinder section (20) at the position with the smallest length along the normal direction of the leading edge impact plate.

7. The cooling structure for the leading edge of a turbine blade according to claim 5, characterized in that: The axis of the cylindrical section (10) coincides with the axis of the conical barrel section (20).

8. The cooling structure for the leading edge of a turbine blade according to claim 5, characterized in that: The length ratio of the cylindrical section (10) to the conical section (20) is 0.5 to 1.

5.

9. The cooling structure for the leading edge of a turbine blade according to claim 5, characterized in that: The length of the cylindrical section (10) is 2 to 4 times the diameter of the cylindrical section (10).

10. The cooling structure for the leading edge of a turbine blade according to claim 1, characterized in that: The air film hole structure is a multi-row and multi-column air film hole structure, and the number of leading edge air film holes (7) in each row arranged along the blade height direction is equal to the sum of the number of the first oblique impact holes (3) and the second oblique impact holes (4).

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