An aircraft engine high pressure turbine blade cascade end wall cooling structure

Abstract

The present application belongs to the field of design of end wall cooling technology for high-pressure turbine blade cascades of aircraft engines, and specifically relates to an end wall cooling structure for high-pressure turbine blade cascades of aircraft engines, comprising: a cooling air inlet hole, which is opened on the outer wall of the end wall; a cooling cavity, which is opened inside the end wall and connected to the cooling air inlet hole; a plurality of air film holes, which are opened on the inner wall of the end wall and connected to the cooling air inlet hole, wherein the air film holes located in the leading edge blade back area have a plurality of pairs of punching holes, and the outflow angles of each pair of punching holes are offset in the circumferential direction of the blade cascade; the air film holes located in the suction surface shoulder have an outflow angle that deviates from the mainstream direction of the high-temperature combustion gas in the blade cascade channel; the air film holes located in the pressure surface corner area have an outflow angle that deviates from the mainstream direction of the high-temperature combustion gas in the blade cascade channel; the air film holes located in the trailing edge corner area have an outflow angle that deviates from the mainstream direction of the high-temperature combustion gas in the blade cascade channel.

Description

Technical Field

[0001] The present application belongs to the field of aircraft engine high-pressure turbine blade end wall cooling technology design, and specifically relates to an aircraft engine high-pressure turbine blade end wall cooling structure. Background Art

[0002] With the increase of aircraft engine thrust, the inlet temperature of the high-pressure turbine increases greatly. The inner wall of the high-pressure turbine blade end is directly exposed to the high-temperature combustion gas and is subjected to extremely high temperature load. In order to increase its service life, a corresponding cooling structure is designed.

[0003] The cooling structure of the end wall of the current aircraft engine high-pressure turbine blade cascade is mainly designed to have a cooling air inlet hole on the outer wall of the end wall, a cooling cavity connected to the cooling air inlet hole inside the end wall, and a plurality of air film holes connected to the cooling cavity on the inner wall of the end wall. Most of the air film holes are evenly distributed on the inner wall of the end wall. The cooling air can enter the cooling cavity from the cooling air inlet hole, and then flow out from the air film holes to form an air film on the inner wall surface of the end wall, thereby realizing the cooling of the inner wall of the end wall and protecting the end wall from erosion by high-temperature combustion gas. This technical solution has the following defects:

[0004] Affected by the structure of the high-pressure turbine cascade of an aircraft engine and the flow characteristics of its high-temperature gas, there are four potential high-temperature heat transfer areas on the end wall, such as Figure 1 As shown, they are:

[0005] The back area of ​​the leading edge blade is impacted by the high-temperature combustion gas sucked by the leading edge horseshoe cavity, which enhances the heat exchange in this area;

[0006] The shoulder of the suction surface is located between the separation line of the suction surface branch horseshoe vortex and the separation line of the blade channel vortex. There is a complex vortex system, and the suction surface profile has an accelerating effect on the high-temperature combustion gas, which enhances the heat exchange in this area.

[0007] The pressure surface angle area, where the high-temperature gas velocity is high and the boundary layer is thin, causes the heat transfer in this area to be enhanced;

[0008] The trailing edge corner area has a high velocity of high-temperature gas and a high degree of turbulence, which enhances the heat exchange in this area.

[0009] The above four potential high-temperature heat transfer areas are limited by the coating process, and the thermal barrier coating on them is very easy to peel off. After the thermal barrier coating peels off, when the air film holes are evenly distributed on the inner wall of the end wall, local high temperature is likely to occur, causing oxidation of the end wall substrate and ablation.

[0010] This application is proposed in view of the above-mentioned technical defects.

[0011] It should be noted that the disclosure of the above background technology content is only used to assist in understanding the inventive concept and technical solution of the present invention, and it does not necessarily belong to the prior art of this patent application. In the absence of clear evidence that the above content has been disclosed on the filing date of this application, the above background technology should not be used to evaluate the novelty and creativity of the present application. Summary of the invention

[0012] The purpose of the present application is to provide an aircraft engine high-pressure turbine blade end wall cooling structure to overcome or alleviate at least one of the known technical deficiencies.

[0013] The technical solution of this application is:

[0014] An aircraft engine high-pressure turbine blade cascade end wall cooling structure, comprising:

[0015] A cooling air inlet hole is provided on the outer wall of the end wall;

[0016] A cooling cavity is provided inside the end wall and communicated with the cooling air inlet hole;

[0017] A plurality of air film holes are opened on the inner wall of the end wall and communicated with the cooling air inlet holes, wherein:

[0018] The air film holes located in the back area of ​​the leading edge blade have multiple pairs of punching holes, and the outflow angles of each pair of punching holes are opposite in the circumferential direction of the blade cascade;

[0019] The air film holes located on the shoulder of the suction surface have an outflow angle that deviates from the mainstream direction of the high-temperature combustion gas in the cascade channel;

[0020] The outflow angle of the air film holes located in the pressure surface angle area deviates from the mainstream direction of the high-temperature combustion gas in the cascade channel;

[0021] The outflow angle of the air film holes located in the trailing edge corner area deviates from the mainstream direction of the high-temperature combustion gas in the blade channel.

[0022] According to at least one embodiment of the present application, in the above-mentioned aircraft engine high-pressure turbine blade cascade end wall cooling structure, each pair of punch holes located in the leading edge blade back region has a punch angle of 30° in the circumferential direction of the blade cascade;

[0023] The pairs of punch holes located in the back area of ​​the leading edge blade are staggered in the axial direction of the blade grid.

[0024] According to at least one embodiment of the present application, in the above-mentioned aircraft engine high-pressure turbine blade end wall cooling structure, the air film holes located on the shoulder of the suction surface have an outflow angle that deviates from the mainstream direction of the high-temperature combustion gas in the blade channel by 20° to 45° and faces away from the back of the corresponding blade.

[0025] According to at least one embodiment of the present application, in the above-mentioned aircraft engine high-pressure turbine blade end wall cooling structure, the air film holes located in the pressure surface angle zone have an outflow angle that deviates from the mainstream direction of the high-temperature combustion gas in the blade channel by 10° to 30°, and is directed toward the blade basin of the corresponding blade.

[0026] According to at least one embodiment of the present application, in the above-mentioned aircraft engine high-pressure turbine blade end wall cooling structure, the air film holes located in the trailing edge corner area have an outflow angle that deviates from the mainstream direction of the high-temperature combustion gas in the blade channel by 10° to 30°, toward the trailing edge of the corresponding blade.

[0027] According to at least one embodiment of the present application, in the above-mentioned aircraft engine high-pressure turbine blade end wall cooling structure, the density of the air film holes located in the leading edge blade back area, suction surface shoulder, pressure surface corner area, and trailing edge corner area is relatively large. BRIEF DESCRIPTION OF THE DRAWINGS

[0028] Figure 1 It is a schematic diagram of the cooling structure of the end wall of the high-pressure turbine blade cascade of an existing aircraft engine;

[0029] Figure 2 It is a schematic diagram of the cooling structure of the end wall of the high-pressure turbine blade cascade of an aircraft engine provided in an embodiment of the present application.

[0030] In order to better illustrate this embodiment, some parts of the drawings may be omitted, enlarged or reduced, and do not represent the size of the actual product. In addition, the drawings are only used for illustrative purposes and cannot be understood as limitations on this patent. DETAILED DESCRIPTION

[0031] In order to make the technical solution and advantages of the present application clearer, the technical solution of the present application will be further described in detail in detail and in detail in conjunction with the accompanying drawings. It can be understood that the specific embodiments described here are only partial embodiments of the present application, which are only used to explain the present application, not to limit the present application. It should be noted that, for the convenience of description, only the parts related to the present application are shown in the accompanying drawings, and other related parts can refer to the general design. In the absence of conflict, the embodiments in the present application and the technical features in the embodiments can be combined with each other to obtain new embodiments.

[0032] In addition, unless otherwise defined, the technical terms or scientific terms used in the description of this application shall be the common meanings understood by the general technicians in the field to which this application belongs. The words "upper", "lower", "left", "right", "center", "vertical", "horizontal", "inside", "outside" and other words indicating orientation used in the description of this application are only used to indicate the relative direction or positional relationship, and do not imply that the device or element must have a specific orientation, be constructed and operated in a specific orientation. When the absolute position of the described object changes, its relative positional relationship may also change accordingly, so it cannot be understood as a limitation on this application. The "first", "second", "third" and similar terms used in the description of this application are only used for descriptive purposes to distinguish different components, and cannot be understood as indicating or implying relative importance. The similar words "one", "one" or "the" used in the description of this application should not be understood as an absolute limitation on quantity, but should be understood as the existence of at least one. The similar words "including" or "comprising" used in the description of this application mean that the elements or objects appearing in front of the word include the elements or objects listed after the word and their equivalents, without excluding other elements or objects.

[0033] In addition, it should be noted that, unless otherwise clearly specified and limited, the words "installed", "connected", "connected" and similar terms used in the description of this application should be understood in a broad sense. For example, the connection can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection or an electrical connection; it can be a direct connection, an indirect connection through an intermediate medium, or the internal connection of two components. Technical personnel in the field can understand their specific meanings in this application according to the specific circumstances.

[0034] The following is combined with Figure 1 to Figure 2 This application is described in further detail.

[0035] An aircraft engine high-pressure turbine blade cascade end wall cooling structure, comprising:

[0036] A cooling air inlet hole is provided on the outer wall of the end wall;

[0037] A cooling cavity is provided inside the end wall and communicated with the cooling air inlet hole;

[0038] A plurality of air film holes are opened on the inner wall of the end wall and communicated with the cooling air inlet holes, wherein:

[0039] The air film holes located in the back area of ​​the leading edge blade have multiple pairs of punching holes, and the outflow angles of each pair of punching holes are opposite in the circumferential direction of the blade cascade;

[0040] The air film holes located on the shoulder of the suction surface have an outflow angle that deviates from the mainstream direction of the high-temperature combustion gas in the cascade channel;

[0041] The outflow angle of the air film holes located in the pressure surface angle area deviates from the mainstream direction of the high-temperature combustion gas in the cascade channel;

[0042] The outflow angle of the air film holes located in the trailing edge corner area deviates from the mainstream direction of the high-temperature combustion gas in the blade channel.

[0043] As for the cooling structure of the end wall of the high-pressure turbine blade cascade of an aircraft engine disclosed in the above-mentioned embodiment, it can be understood by technicians in the field that the air film holes designed to be located in the leading edge blade back area have multiple pairs of counter-punching holes, and the cooling gas flowing out of each pair of counter-punching air film holes counter-punches in the circumferential direction of the blade cascade, which can effectively weaken the leading edge horseshoe vortex and can better adhere to the inner wall of the end wall, thereby achieving a better cooling effect on the leading edge blade back area. In addition, the air film holes designed to be located in the suction surface shoulder, the suction surface shoulder, and the trailing edge corner area have an outflow angle that deviates from the mainstream direction of the high-temperature combustion gas in the blade cascade channel, so as to avoid the cooling gas flowing out of the air film holes directly flowing away with the high-temperature combustion gas in the blade cascade channel, inhibit the development of the vortex system, and can better adhere to the inner wall of the end wall, thereby achieving a better cooling effect on the suction surface shoulder, the suction surface shoulder, and the trailing edge corner area.

[0044] Regarding the aircraft engine high-pressure turbine blade end wall cooling structure disclosed in the above-mentioned embodiment, technical personnel in the field can also understand that, in order to address the causes of the four potential high-temperature heat transfer areas on the end wall, the air film holes in the leading edge blade back area, the suction surface shoulder, the suction surface shoulder, and the trailing edge corner area are differentially designed, or only some of the air film holes are differentiated, so as to achieve better cooling effects on the leading edge blade back area, the suction surface shoulder, the suction surface shoulder, and the trailing edge corner area. In the event that the thermal barrier coating thereon peels off, local high temperatures and substrate oxidation can be avoided, thereby protecting the end wall from being ablated.

[0045] In some optional embodiments, in the above-mentioned aircraft engine high-pressure turbine blade cascade end wall cooling structure, each pair of punch holes located in the leading edge blade back region has an outflow angle of 30° in the circumferential direction of the blade cascade and can be 30° to 60° in the axial direction;

[0046] The pairs of punch holes located in the back area of ​​the leading edge are staggered in the axial direction of the blade grid to supplement the interval cooling between the pairs of punch holes.

[0047] In some optional embodiments, in the above-mentioned aircraft engine high-pressure turbine blade end wall cooling structure, the air film holes located on the shoulder of the suction surface have an outflow angle that deviates from the mainstream direction of the high-temperature combustion gas in the blade channel by 20° to 45°, and faces away from the back of the corresponding blade, that is, the edge of the inclined end wall, so as to ensure the cooling effect on this part on the suction side.

[0048] In some optional embodiments, in the above-mentioned aircraft engine high-pressure turbine blade end wall cooling structure, the air film holes located in the pressure surface angle zone have an outflow angle that deviates from the mainstream direction of the high-temperature combustion gas in the blade channel by 10° to 30°, toward the blade basin of the corresponding blade, and selects the smallest possible angle in the circumferential direction of the blade, which can be specifically 30°, to ensure the cooling effect on this part on the pressure side.

[0049] In some optional embodiments, in the above-mentioned aircraft engine high-pressure turbine blade end wall cooling structure, the air film holes located in the trailing edge corner area have an outflow angle that deviates from the mainstream direction of the high-temperature combustion gas in the blade channel by 10° to 30°, toward the trailing edge of the corresponding blade, so as to suppress the development of the vortex system in the trailing edge corner area.

[0050] In some optional embodiments, in the above-mentioned aircraft engine high-pressure turbine blade end wall cooling structure, the density of the air film holes located in the leading edge blade back area, suction surface shoulder, pressure surface corner area, and trailing edge corner area is relatively large to enhance the cooling effect on the leading edge blade back area, suction surface shoulder, pressure surface corner area, and trailing edge corner area. In the event that the thermal barrier coating thereon peels off, local high temperature and substrate oxidation can be avoided, thereby protecting the end wall from being ablated.

[0051] The various embodiments in the specification are described in a progressive manner, and each embodiment focuses on the differences from other embodiments. The same or similar parts between the various embodiments can be referenced to each other.

[0052] So far, the technical solution of the present application has been described in conjunction with the preferred embodiments shown in the accompanying drawings. Those skilled in the art should understand that the scope of protection of the present application is obviously not limited to these specific embodiments. Without departing from the principles of the present application, those skilled in the art can make equivalent changes or substitutions to the relevant technical features, and the technical solutions after these changes or substitutions will fall within the scope of protection of the present application.

Claims

1. An aircraft engine high pressure turbine blade end wall cooling structure, characterized in that: include: A cooling air inlet hole is provided on the outer wall of the end wall; A cooling cavity is provided inside the end wall and communicated with the cooling air inlet hole; A plurality of air film holes are opened on the inner wall of the end wall and communicated with the cooling air inlet holes, wherein: The air film holes located in the back area of ​​the leading edge blade have multiple pairs of punching holes, and the outflow angles of each pair of punching holes are opposite in the circumferential direction of the blade cascade; The air film holes located on the shoulder of the suction surface have an outflow angle that deviates from the mainstream direction of the high-temperature combustion gas in the cascade channel; The outflow angle of the air film holes located in the pressure surface angle area deviates from the mainstream direction of the high-temperature combustion gas in the cascade channel; The outflow angle of the air film holes located in the trailing edge corner area deviates from the mainstream direction of the high-temperature combustion gas in the blade channel.

2. The aircraft engine high pressure turbine blade end wall cooling structure according to claim 1, characterized in that: The outflow angle of each pair of punch holes located in the back area of ​​the leading edge blade is 30° in the circumferential direction of the blade cascade; The pairs of punch holes located in the back area of ​​the leading edge blade are staggered in the axial direction of the blade grid.

3. The aircraft engine high pressure turbine blade end wall cooling structure according to claim 1, characterized in that: The outflow angle of the air film holes located on the shoulder of the suction surface deviates from the mainstream direction of the high-temperature combustion gas in the blade channel by 20° to 45°, and faces away from the back of the corresponding blade.

4. The aircraft engine high pressure turbine blade end wall cooling structure according to claim 1, characterized in that: The outflow angle of the air film holes located in the pressure surface angle zone deviates from the mainstream direction of the high-temperature combustion gas in the blade cascade channel by 10° to 30°, and is directed toward the blade basin of the corresponding blade.

5. The aircraft engine high pressure turbine blade end wall cooling structure according to claim 1, characterized in that: The outflow angle of the air film holes located in the trailing edge corner area deviates from the mainstream direction of the high-temperature combustion gas in the blade channel by 10° to 30°, and is directed toward the trailing edge of the corresponding blade.

6. The aircraft engine high pressure turbine blade cascade end wall cooling structure according to claim 1, characterized in that: The density of the air film holes located in the leading edge blade back area, suction surface shoulder, pressure surface corner area, and trailing edge corner area is relatively large.

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