Double-wall blade cooling structure, blade, aircraft engine and gas turbine

Abstract

The present invention provides a double-wall blade cooling structure, a blade, an aircraft engine and a gas turbine, wherein a cold air channel is arranged between the inner wall and the outer wall of the blade, and an air supply hole arranged on the inner wall of the blade introduces the airflow in the cold air cavity of the blade into the cold air channel, and the convection heat exchange effect of the flow impact hole and the spoiler column in the cold air channel is used to improve the internal heat exchange capacity, and the cooling air flows out through the air film holes on the outer wall of the blade after impacting the outer wall of the blade, thereby realizing the cooling of the outer wall of the blade, improving the cooling effect of the turbine blade, and increasing the service life of the turbine blade.

Description

Technical Field

[0001] The invention relates to the field of turbines, and in particular to a double-wall blade cooling structure, a blade, an aircraft engine and a gas turbine. Background Art

[0002] In order to pursue better overall thermal efficiency, according to the Brayton cycle principle, the temperature before the turbine of aircraft engines and gas turbines has been continuously increased, and the gas temperature at the combustion chamber outlet has far exceeded the allowable temperature of the metal material of the turbine blades. Since the first proposal of cooling blades in the 1940s and 1950s, turbine blade cooling technology has enabled turbine blades to operate in a temperature environment that exceeds the long-term working temperature of blade metal.

[0003] Blade cooling technology mainly includes internal enhanced heat exchange cooling and external film cooling. After a long period of development and accumulation, the efficiency of internal enhanced heat exchange and external film cooling has reached a very high level. Conventional cooling methods are difficult to support the further increase in the temperature before the turbine of future aircraft engines and gas turbines. As a highly efficient and compact cooling structure, double-wall cooling has great potential for application in the cooling of current aircraft engine or gas turbine turbine blades. In recent years, due to the advancement of processing technology, experimental capabilities and computer numerical simulation capabilities, double-wall cooling structures have been successfully applied in the development of domestic and foreign aircraft engines. In the future, by improving the double-wall design method, the overall performance and life of aircraft engines or gas turbines can be greatly improved. Summary of the invention

[0004] In view of this, the embodiments of this specification provide a double-wall blade cooling structure, a blade, an aircraft engine and a gas turbine, which are used to improve the cooling effect of the turbine blades and increase the life of the turbine blades.

[0005] The embodiments of this specification provide the following technical solutions: A double-wall blade cooling structure, characterized in that it includes: a double-wall cooling structure body, provided with an air supply hole and a cold air channel located inside, the air supply hole is arranged at the inner wall of the double-wall cooling structure body, and the two ends of the air supply hole are respectively connected to the cold air channel and the outside of the double-wall cooling structure body; a flow impact plate, arranged inside the cold air channel, the flow impact plate is provided with a flow impact hole, and the axial direction of the flow impact hole is parallel to the gas flow direction in the cold air channel; a spoiler column, arranged in the cold air channel and located on the side of the flow impact plate away from the air supply hole; an air film hole, arranged at the outer wall of the double-wall cooling structure body, one end of the air film hole is connected to the cold air channel at the spoiler column, and the other end of the air film hole is connected to the outside of the double-wall cooling structure body.

[0006] Furthermore, the cross-sectional shapes of the flow-toward impact hole include circular, elliptical and rectangular.

[0007] Furthermore, the cross-sectional shapes of the spoiler column include circular, elliptical, rectangular, teardrop-shaped and triangular.

[0008] Furthermore, the cross-sectional shape of the air film hole is circular, wedge-shaped or slit-shaped.

[0009] The embodiment of the present specification also provides a blade, wherein the blade includes the double-wall blade cooling structure.

[0010] Furthermore, the blade includes a blade body, a cold air cavity is arranged inside the blade body, the double-wall blade cooling structure is arranged on the pressure surface and / or suction surface of the blade body, and the cold air cavity is connected to the double-wall cooling structure body through an air supply hole.

[0011] Furthermore, there are a plurality of air supply holes, and the plurality of air supply holes are arranged at intervals along the height direction of the blade body.

[0012] Furthermore, the flow impact plate is provided with a plurality of flow impact holes, the blade further comprises a plurality of spoiler columns, and the plurality of flow impact holes are provided in a one-to-one correspondence with the plurality of spoiler columns.

[0013] Furthermore, a plurality of spoiler columns are arranged at intervals along the height direction of the blade body to form a spoiler column row.

[0014] Furthermore, the spoiler column rows are multiple rows, and the multiple spoiler columns in adjacent spoiler column rows are arranged in a sequential row or a staggered row.

[0015] An embodiment of the present specification also provides an aircraft engine, which includes blades with a double-wall cooling structure.

[0016] An embodiment of the present specification also provides a gas turbine, wherein the gas turbine includes blades with a double-wall cooling structure.

[0017] Compared with the prior art, at least one of the above-mentioned technical solutions adopted in the embodiments of this specification can achieve at least the following beneficial effects: by providing a double-wall blade cooling structure, a cold air channel is set between the inner wall of the blade and the outer wall of the blade, and the air supply holes set on the inner wall of the blade introduce the airflow in the cold air cavity into the cold air channel, and the internal heat exchange capacity is improved through the flow impact holes and spoiler columns in the cold air channel, thereby achieving the technical effect of improving the cooling effect of the turbine blades and increasing the service life of the turbine blades. BRIEF DESCRIPTION OF THE DRAWINGS

[0018] 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.

[0019] Figure 1 is a schematic diagram of a double-wall blade cooling structure provided by an embodiment of the present invention;

[0020] Figure 2 is a cross-sectional view of a double-wall blade cooling structure provided by an embodiment of the present invention;

[0021] Figure 3 is a schematic diagram of a blade structure provided by an embodiment of the present invention;

[0022] Figure 4 It is a schematic diagram of a blade profile provided by an embodiment of the present invention.

[0023] The reference numerals in the figure are: 1, blade structure; 11, blade body; 12, edge plate; 13, root; 14, cold air cavity; 110, leading edge; 111, pressure surface; 112, suction surface; 113, trailing edge; 200, double-wall cooling structure body; 201, air supply hole; 202, cold air channel; 203, flow impact hole; 204, spoiler column; 205, air film hole; 206, blade body inner wall; 207, blade body outer wall; 208, side wall. DETAILED DESCRIPTION

[0024] The embodiments of the present application are described in detail below with reference to the accompanying drawings.

[0025] The following describes the implementation methods of the present application through specific examples, and those skilled in the art can easily understand other advantages and effects of the present application from the contents 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, in the absence of conflict, the following embodiments and the features in the embodiments can be combined with each other. 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.

[0026] like Figure 1As shown, an embodiment of the present invention provides a double-wall blade cooling structure, including: a double-wall cooling structure body 200, provided with an air supply hole 201 and a cold air channel 202 located inside, the air supply hole 201 is arranged at the inner wall of the double-wall cooling structure body 200, and the two ends of the air supply hole 201 are respectively connected to the cold air channel 202 and the outer side of the double-wall cooling structure body 200; a flow impact plate, arranged inside the cold air channel 202, and the flow impact plate is provided with a flow impact hole 203, and the axial direction of the flow impact hole 203 is parallel to the gas flow direction in the cold air channel 202; a spoiler column 204, arranged in the cold air channel 202 and located on the side of the flow impact plate away from the air supply hole 201; an air film hole 205, arranged at the outer wall of the double-wall cooling structure body 200, one end of the air film hole 205 is connected to the cold air channel 202 at the spoiler column 204, and the other end of the air film hole 205 is connected to the outer side of the double-wall cooling structure body 200.

[0027] In this embodiment, the air flow is introduced into the cold air channel 202 through the air supply hole 201, and the internal heat exchange capacity is improved through the flow impact hole 203 and the spoiler column 204 in the cold air channel 202, thereby achieving the purpose of improving the cooling effect of the turbine blades and increasing the life of the turbine blades.

[0028] Specifically, Figure 1 and Figure 2 As shown, the double-wall cooling structure body 200 is composed of an air supply hole 201 , a cold air channel 202 , a flow impact hole 203 , a spoiler column 204 and an air film hole 205 . Figure 2 for Figure 1The internal cross-sectional view of the double-wall cooling structure body 200, i.e., the cross-sectional view of the BB plane. The cold air channel is composed of the inner wall 206 of the blade body, the outer wall 207 of the blade body, and the side wall 208. There are multiple air supply holes 201, which are arranged on the inner wall 206 of the blade body. The cooling air in the cold air channel 202 flows in through the air supply holes 201 arranged on the inner wall 206 of the blade body, flows out through the air film holes 205 arranged on the outer wall 207 of the blade body, and impacts the outer wall 207 of the blade body of the turbine blade, which has the effect of enhancing the heat exchange of the outer wall 207 of the blade body. The multiple flow impact holes 203 on the flow impact plate are arranged in the cold air channel 202 opposite to the multiple spoiler columns 204, and the spoiler columns 204 are arranged on the side away from the air supply holes 201. The cooling air first flows through the impact hole 203 in the cold air channel 202, and then impacts the downstream spoiler column 204, thereby strengthening the heat exchange capacity of the spoiler column 204, so that the heat exchange of the spoiler column 204 changes from conventional airflow sweeping heat exchange to impact sweeping heat exchange, and the heat conductivity of the spoiler column 204 is increased. The air film hole 205 is a film cooling hole structure. After the air flow flows through the air film hole 205, a layer of air flow with a lower temperature is formed on the outer surface of the outer wall 207 of the turbine blade body, which can play a role in preventing high-temperature combustion gas from eroding the turbine blade body. The air film hole 205 is arranged on the outer wall 207 of the blade body. After the cold air flows through the spoiler column 204, it flows out through the air film hole 205, which has the effect of air film cooling on the outer wall 208 of the blade body.

[0029] Preferably, the diameter of the air supply holes 201 is 0.5 mm to 2 mm, and the ratio of the hole spacing between adjacent air supply holes 201 to the diameter of each air supply hole 201 is in the range of 3 to 15.

[0030] Optionally, the cross-sectional shape of the air supply hole 201 includes circular, elliptical, rectangular and other shapes.

[0031] Optionally, the cross-sectional shape of the flow impact hole 203 includes circular, elliptical, rectangular and other shapes.

[0032] Optionally, the cross-sectional shape of the spoiler column 204 includes circular, elliptical, rectangular, teardrop-shaped, triangular and the like.

[0033] Optionally, the cross-sectional shape of the air film hole 205 includes circular, wedge-shaped, and slit-shaped shapes, and the corresponding hole types include cylindrical holes, expanded holes, wedge-shaped holes, slit-shaped holes, etc. Among them, the expanded hole includes a small end face and a large end face, and the shapes of the small end face and the large end face can be any shape, and the cross-sectional area of ​​the expanded hole gradually increases along the extension direction of the small end face to the large end face.

[0034] Preferably, when the cross-sectional shape of the air film hole 205 is circular, the diameter of the air film hole 205 ranges from 0.3 mm to 1.5 mm.

[0035] like Figure 3 and Figure 4 As shown, Figure 4 is the blade profile diagram, that is Figure 3 Cross-sectional view of the AA plane in the middle. An embodiment of the present invention further provides a blade structure 1, wherein the blade structure 1 includes the double-wall cooling structure body 200. The blade structure 1 includes: a blade body 11, an edge plate 12 and a root 13. The turbine rotor blade structure 1 has a channel for cooling air to pass through, and the cold air enters the interior of the turbine blade blade body 11 through the channel on the root 13. The blade body 11 has a cold air cavity 14 inside, and the blade body 11 includes a leading edge 110, a pressure surface 111, a suction surface 112 and a trailing edge 113. Among them, the leading edge 110 and the trailing edge 113 are arranged opposite to each other, and the pressure surface 111 and the suction surface 112 are connected through the leading edge 110 and the trailing edge 113. The double-wall cooling structure body 200 proposed in the present invention is arranged on the pressure surface 111 and / or the suction surface 112 of the blade body 11, and the cold air cavity 14 is connected to the double-wall cooling structure body 200 through the air supply hole 201.

[0036] Specifically, the turbine blade is a heat exchanger device having an air passage inside. The blade body 11 is an airfoil, a gas flow passage is formed inside the edge plate 12, and the root 13 has a wavy profile. The double-wall cooling structure body 200 is applied to the blade body 11, and the cold air passage 202 is formed in the thickness direction of the blade body 11, thereby forming a blade body inner wall 206 and a blade body outer wall 207. The cooling air in the cold air passage 202 flows in through the air supply hole 201 provided on the blade body inner wall 206, and flows out through the air film hole 205 provided on the blade body outer wall 207, thereby cooling the outer wall 207 of the turbine blade body.

[0037] Preferably, the thickness of the blade inner wall 206 is 0.5 mm to 2 mm, the thickness of the blade outer wall 207 is 0.5 mm to 2 mm, and the height of the cold air passage 202 along the blade wall thickness direction is 0.5 mm to 2 mm.

[0038] Preferably, the plurality of air supply holes 201 on the inner wall 206 of the blade body are arranged at intervals along the height direction of the blade body 11, including uniform intervals and non-uniform intervals, and the arrangement method is determined according to the heat load of the outer wall of the blade body. By changing the shape and size of the air supply holes 201, the impact heat exchange intensity can be effectively adjusted, and by adjusting the layout of the air supply holes 201 along the height direction of the blade body 11, the impact heat exchange effect of different parts of the blade body can be improved.

[0039] Preferably, the flow impact plate is provided with a plurality of flow impact holes 203 , and the blade body 11 further comprises a plurality of spoiler columns 204 , and the plurality of flow impact holes 203 and the plurality of spoiler columns 204 are provided in a one-to-one correspondence.

[0040] Preferably, a plurality of spoiler columns 204 are arranged at intervals along the height direction of the blade body 11 to form spoiler column rows.

[0041] Preferably, the spoiler column rows are multiple rows, and the multiple spoiler columns 204 in adjacent spoiler column rows are arranged in a sequential row or a staggered row.

[0042] Specifically, the spoiler columns 204 flowing downstream of the impact hole 203 can be designed as double rows to further enhance the heat exchange efficiency. In addition, the layout of the two rows of spoiler columns 204 can be arranged in a staggered or sequential manner to improve the heat conductivity of the downstream spoiler columns 204.

[0043] Optionally, the air film holes 205 may be designed as cylindrical holes, expansion holes or slot holes, and different hole types may be applied at different positions of the blade body 11 to improve the overall cooling effect.

[0044] Preferably, by improving the exhaust mode of the cold air in the cold air passage 202 from exhaust through the air outlet holes to exhaust through the trailing edge slits, the trailing edge slit structure can be applied to the suction surface 112 of the turbine blade body 11 near the trailing edge 113 .

[0045] Further, in the embodiment of the present invention, the specific cold air flow path and the scheme for improving the cooling effect are as follows: the cold air flows into the inside of the blade through the cold air channel inside the root 13 of the turbine rotor blade, and supplies the air source required for cooling the double-wall cooling structure body 200. The cold air reaches the cold air cavity 14 of the blade body 11. Since the cold air pressure in the cold air cavity 14 is relatively high, driven by the pressure, the cold air passes through the air supply hole 201 on the inner wall 206 of the blade body and impacts the outer wall 207 of the blade body, which has the effect of strengthening the heat exchange of the outer wall 207 of the blade body. By changing the shape and size of the air supply hole 201, the impact heat exchange intensity can be effectively adjusted. By adjusting the layout of the air supply hole 201 along the height direction of the blade body 11, the impact heat exchange effect of different parts of the blade body can be improved. After the impact, the cold air reaches the cold air channel 202 and flows along the blade flow direction. In the cold air channel 202, the airflow first passes through the flow impact hole 203 and impacts the downstream spoiler column 204, which strengthens the heat exchange capacity of the spoiler column 204. The heat exchange of the spoiler column 204 is changed from conventional airflow sweep heat exchange to impact sweep heat exchange, which increases the thermal conductivity of the spoiler column 204. The spoiler column 204 flowing downstream of the impact hole 203 can be designed as a double row to further enhance the heat exchange efficiency. In addition, the layout of the two rows of spoiler columns 204 can be arranged in a staggered or sequential manner to improve the thermal conductivity of the downstream spoiler column 204. After the cold air flows through the spoiler column 204, it flows out through the air film hole 205, which has the effect of air film cooling on the outer wall 207 of the blade body. The air film hole 205 can be designed as a cylindrical hole, an expansion hole and a slit hole. Different hole types are applied at different positions of the blade body 11 to improve the overall cooling effect.

[0046] In addition, the embodiment of the present invention further provides an aircraft engine and a gas turbine having the above blade structure 1, wherein the blade structure 1 is a blade with a double-wall cooling structure body 200. The novel double-wall cooling structure of the aircraft engine or gas turbine of the present invention improves the double-wall cooling structure layout, greatly improves the internal heat exchange effect, weakens the external air film cooling effect, improves the cooling effect, and reduces the thermal stress problem caused by excessive external temperature gradient, which plays a positive and effective role in the thermal efficiency and life of the aircraft engine or gas turbine.

[0047] In summary, the double-wall blade cooling structure, blade, aircraft engine and gas turbine provided in the embodiments of the present application can achieve at least the following technical effects compared with the prior art:

[0048] (1) By introducing flow impingement holes, the heat transfer coefficient of internal cooling is improved;

[0049] (2) By increasing the internal heat transfer coefficient, the cold air temperature is increased and the temperature of the external air film holes is reduced, thereby reducing problems such as excessive thermal stress caused by the air film holes.

[0050] The above is only a specific implementation of the present application, but the protection scope of the present application is not limited thereto. Any changes or substitutions that can be easily thought of by a person skilled in the art within the technical scope disclosed in the present application should be included in the protection scope of the present application. Therefore, the protection scope of the present application shall be based on the protection scope of the claims.

Claims

1. A double-wall blade cooling structure, characterized in that: include: A double-wall cooling structure body (200) is provided with an air supply hole (201) and a cold air channel (202) located inside, wherein the air supply hole (201) is arranged on the inner wall of the double-wall cooling structure body (200), and two ends of the air supply hole (201) are respectively connected to the cold air channel (202) and the outer side of the double-wall cooling structure body (200); A flow impact plate is arranged inside the cold air channel (202), and a flow impact hole (203) is arranged on the flow impact plate, and the axial direction of the flow impact hole (203) is parallel to the gas flow direction in the cold air channel (202); A spoiler column (204) is arranged in the cold air channel (202) and is located on a side of the flow impact plate away from the air supply hole (201); The air film hole (205) is arranged on the outer wall of the double-wall cooling structure body (200), one end of the air film hole (205) is connected to the cold air channel (202) at the spoiler column (204), and the other end of the air film hole (205) is connected to the outer side of the double-wall cooling structure body (200).

2. A double-wall blade cooling structure according to claim 1, characterized in that: The cross-sectional shapes of the flow-direction impact hole (203) include circular, elliptical and rectangular.

3. The double-wall blade cooling structure according to claim 1, characterized in that: The cross-sectional shape of the spoiler column (204) includes circular, elliptical, rectangular, teardrop-shaped and triangular.

4. A double-wall blade cooling structure according to claim 1, characterized in that: The cross-sectional shape of the air film hole (205) is circular, wedge-shaped or slit-shaped.

5. A blade, comprising a double-wall blade cooling structure, characterized in that: The double-wall blade cooling structure is the double-wall blade cooling structure according to any one of claims 1 to 4.

6. A blade according to claim 5, characterized in that: The blade comprises a blade body (11), a cold air cavity (14) is arranged inside the blade body (11), the double-wall blade cooling structure is arranged on the pressure surface (111) and / or the suction surface (112) of the blade body (11), and the cold air cavity (14) is connected to the double-wall cooling structure body (200) through an air supply hole (201).

7. A blade according to claim 6, characterized in that: There are a plurality of air supply holes (201), and the plurality of air supply holes (201) are arranged at intervals along the height direction of the blade body (11).

8. A blade according to claim 6, characterized in that: The flow impact plate is provided with a plurality of flow impact holes (203), and the blade further comprises a plurality of spoiler columns (204), and the plurality of flow impact holes (203) and the plurality of spoiler columns (204) are provided in a one-to-one correspondence.

9. A blade according to claim 8, characterized in that: A plurality of spoiler columns (204) are arranged at intervals along the height direction of the blade body (11) to form a spoiler column row.

10. A blade according to claim 9, characterized in that: The spoiler column rows are multiple rows, and the multiple spoiler columns (204) in adjacent spoiler column rows are arranged in a sequential row or a staggered row.

11. An aircraft engine, comprising a blade with a double-wall cooling structure, characterized in that: The blade is the blade according to any one of claims 5 to 10.

12. A gas turbine comprising blades with a double-wall cooling structure, characterized in that: The blade is the blade according to any one of claims 5 to 10.

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