AIRFOIL OF GAS TURBINE ENGINE

An airfoil of gas turbine engine is provided, which can improve cooling efficiency at a leading edge of the airfoil even when a location of combustion gas stagnation changes, without having to increase cooling air flowrate. The airfoil of gas turbine engine includes a cooling passage formed inside the airfoil to guide cooling air, and one or more cooling slots in elongated form, extending through the leading edge of the airfoil to a direction of the cooling passage and across a lengthwise direction of the leading edge.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application claims priority from Korean Patent Application No. 10-2013-0156847, filed on Dec. 17, 2013, in the Korean Intellectual Property Office, the disclosure of which is incorporated herein by reference in its entirety.

BACKGROUND

1. Field of the Invention

The invention relates to an airfoil of gas turbine engine, and more particularly, to an airfoil of gas turbine engine having efficient film cooling structure at a leading edge thereof.

2. Description of the Related Art

Gas turbine engine has increasing engine efficiency, as the temperature of the gas entering thereto rises. The temperature at the gas turbine entry keeps increasing to meet continuing demands for higher engine efficiency. Accordingly, most recently-available gas turbine engines have turbine entry temperature that is higher than the melting point. It is thus necessary to provide technology to appropriately cool down the related components to prevent melting or failure.

Referring toFIG. 1A, the related art turbine nozzle vanes12include an inner platform16, an outer platform18, and an airfoil section20extending radially (i.e., transversal direction inFIG. 1A) between the inner and outer platforms16,18to form gas flow passage14, for the cooling purpose. To be specific, the airfoil20includes film cooling holes38in cylindrical configuration to cool the surface of the airfoil20, and the platforms16,18have cylindrical film cooling holes40to cool the surfaces of the platforms.

Accordingly, as illustrated inFIG. 1A, some of the cooling air guided from the plenum regions34,36cools the surfaces of the platforms16,28as the air is exhausted onto the surfaces of the platforms16,28through the film cooling holes40. Referring toFIGS. 1A and 1B, the rest of the cooling air that is guided from the plenum regions34,36is introduced into the airfoil cavity24(seeFIG. 1B) and then exhausted onto the surface of the airfoil20, thus cooling the surface of the airfoil20. Reference numeral22denotes turbine blade,26is pressure wall of the airfoil20, and28is a suction wall of the airfoil20.

Referring toFIG. 2A, the related art turbine blade30includes hollow platform40, and a hollow airfoil42extending radially on the platform40(upward direction inFIG. 2A), for cooling purpose. To be specific, the airfoil42includes cylindrical film cooling holes to cool the surface of the airfoil42, and the platform40has cylindrical film cooling holes113to cool the surface of the platform40.

Accordingly referring toFIG. 2A, some of the cooling air that is guided from the hollow cavity (not illustrated) of the platform40is exhausted onto the surface of the platform40through the film cooling holes113, thus cooling the surface of the platform40. Referring toFIGS. 2A and 2B, the rest of the cooling air that is guided from the hollow cavity (not illustrated) of the platform40is introduced into the cavity (seeFIG. 2B) of the airfoil42and then exhausted onto the surface of the airfoil42through the film cooling holes82, thus cooling the surface of the airfoil42. Reference numeral44denotes a pressure side of the airfoil, and46is a suction side of the airfoil.

However, the related art turbine nozzle vane and turbine blade have the following common shortcomings.

That is, since the film cooling holes, which are in circular shape when viewed from the surface of the leading edge, are in the cylindrical shape that are extending from the surface to the cavity, a generous amount of cooling flow is necessary to avoid overheating between the film cooling holes, which leads into increasing use of cooling air flowrate and deteriorating gas turbine efficiency.

Further, while thermal load frequently occurs at the leading edge as the flow of combustion gas has stagnation, it is difficult to accurately predict the location of stagnation in the designing stage, and since the location of stagnation varies depending on the state in which the engine operates in the case of turbine blade, stagnation of high temperature combustion gas between film cooling holes can cause increasing temperature.

SUMMARY

Exemplary embodiments of the present inventive concept overcome the above disadvantages and other disadvantages not described above. Also, the present inventive concept is not required to overcome the disadvantages described above, and an exemplary embodiment of the present inventive concept may not overcome any of the problems described above.

The invention is proposed to solve the problems as described above, and the objective is to provide an airfoil of gas turbine engine which can increase cooling efficiency at a leading edge of the airfoil even when location of combustion gas stagnation changes, without increasing cooling air flowrate.

In one embodiment, an airfoil of gas turbine engine may include a cooling passage formed inside the airfoil to guide cooling air, and one or more cooling slots in elongated form, extending through the leading edge of the airfoil to a direction of the cooling passage and across a lengthwise direction of the leading edge.

One end of each of the one or more cooling slots may be on a compressive surface of the airfoil, and the other end thereof may be on a suction surface of the airfoil.

Both corners of each of the one or more cooling slots may be rounded to a circular arc shape.

The one or more cooling slots may include a first cooling slot and a second cooling slot, and the film cooling structure may additionally include one or more cooling holes formed on the leading edge between the first and second cooling slots, in a manner of extending through the leading edge in a cylindrical shape to a direction of the cooling passage.

The one or more cooling slots may be slanted with respect to an end of the airfoil.

In one embodiment, the one or more cooling slots may be so slanted that the cooling slots become closer to the end of the airfoil at the cooling passage than at the leading edge.

In another embodiment, the one or more cooling slots may be so slanted that the cooling slots become farther away from the end of the airfoil at the cooling passage than at the leading edge.

The airfoil may be an airfoil of a turbine nozzle vane.

The airfoil may be an airfoil of a turbine blade.

According to various embodiments, since one or more cooling slots in elongated configuration are extended through the leading edge of the airfoil to the direction of the cooling passage, and across the lengthwise direction of the leading edge, compared to the related art cooling holes in cylindrical configuration, the cooling efficiency of the leading edge of the airfoil can be greatly improved even when the location of combustion gas stagnation changes at the leading edge, without having to increase the cooling air flowrate. Further, since the cooling air (cooling flow) is exhausted broadly in the lengthwise direction, wider area can be uniformly cooled without having problem such as flow separation from the surface of the vane (or blade).

DETAILED DESCRIPTION OF EXEMPLARY EMBODIMENTS

Certain exemplary embodiments of the present inventive concept will now be described in greater detail with reference to the accompanying drawings.

FIG. 3is a schematic perspective view of an airfoil of a gas turbine engine having a film cooling structure at a leading edge thereof, according to a first embodiment, andFIG. 4is a cross section view of the airfoil ofFIG. 3, taken on line IV-IV.

According to a first embodiment, the airfoil of gas turbine engine includes a cooling passage210, and one or more cooling slots220.

Referring toFIG. 4, the cooling passage210is formed inside the airfoil200to guide the cooling air from a cavity (or plenum region) (34or36inFIG. 1A) of a platform (16or18inFIG. 1A,40inFIG. 2A) into the one or more cooling slots220. Although not illustrated, the cooling passage210may be divided into a plurality of spaces defined by a plurality of partitions.

Referring toFIGS. 3 and 4, the cooling slots220are passed through the leading edge201of the airfoil200to the direction of the cooling passage210. To be specific, the cooling slots220may be elongated across the lengthwise direction (vertical direction inFIG. 3) of the leading edge. The expression “elongated” refers to a certain shape of the hole that is distinguished from the cylindrical shape of the related art with circular cross section, which may be elliptical shape in which the length of the hole is extended to a certain direction, or slot shape with extended length.

Accordingly, since the cooling passage210is passed through the leading edge201of the airfoil200to the direction of the cooling passage210, and one or more elongated cooling slots220are formed across the lengthwise direction of the leading edge201, compared to the related art cooling holes (38inFIG. 1A,82inFIG. 2A) with cylindrical shape, it is possible to increase cooling efficiency of the leading edge201of the airfoil200even when the location of combustion gas stagnation varies, without having to increase the cooling air flowrate.

Further, referring toFIGS. 3 and 4, one end of the cooling slot220may be placed on a compressive surface202of the airfoil200, while the other end thereof may be placed on a suction surface203of the airfoil200. Accordingly, since the cooling slots220are formed across the suction surface203, it is possible to increase cooling efficiency of the leading edge201of the airfoil200even when the location of combustion gas stagnation further changes.

Further, referring toFIG. 3, both corners220aof the cooling slot220may be rounded in the shape of circular arc. Accordingly, it is possible to prevent stress concentration from generating on both corners of the cooling slot220.

The airfoil220may be the airfoil of the turbine nozzle vane (seeFIG. 1A) or the airfoil of the turbine blade (seeFIG. 2A).

The airfoil of the gas turbine engine according to a second embodiment will be explained below, with reference toFIG. 5.

FIG. 5is a schematic perspective view of an airfoil of a gas turbine engine having a film cooling structure at a leading edge thereof, according to a second embodiment.

Referring toFIG. 5, the airfoil of gas turbine engine according to the second embodiment is similar to that according to the first embodiment, except for one or more cooling holes230additionally formed between the cooling slots221,222of the leading edge. Accordingly, the cooling holes230formed between the cooling slots221,222will be explained in detail, while the description of the like elements will be referenced to the explanation given above. Additionally, the same reference numerals and names will be given to the similar or same elements.

The one or more cooling slots220may include first and second cooling slots221,222. For the explanation of the first and second cooling slots221,222, the explanation of the cooling slots220given above in the first embodiment will be referenced.

The one or more cooling holes230may be formed in the leading edge201, i.e., between the first and second cooling slots221,222. For example, the cooling holes230may be formed in a known manner so that the cooling holes230may be cylindrical holes that are passed through to the direction of the cooling passage210.

Since one or more elongated cooling slots220are provided, it is possible to increase cooling efficiency of the leading edge201of the airfoil200even when the location of combustion gas stagnation at the leading edge201varies, without having to increase the cooling air flowrate. Further, since one or more cylindrically-extending cooling holes230are provided, the same cooling function as that of the related art cooling holes (38inFIG. 1A,82inFIG. 2A) may be additionally used.

The airfoil of gas turbine engine according to a third embodiment will be explained below, with reference toFIG. 6.

FIG. 6is a schematic longitudinal cross-section view of a film cooling structure at a leading edge of an airfoil of a gas turbine engine, according to a third embodiment.

Referring toFIG. 6, the film cooling structure at a leading edge of an airfoil of gas turbine engine according to the third embodiment is similar to that according to the first embodiment explained above, except for the cooling slots3220that have compound angle. Accordingly, the cooling slots3220with compound angle will be explained in detail below, while description of the other similar or same elements is referenced to the explanation given above.

The cooling slots3220may be slanted with respect to an end3204of the airfoil3200. To be specific, the cooling slots3220may be slanted at such a compound angle that the cooling slots3220become closer to the end3204of the airfoil3200at the cooling passage3210than at the leading edge3201. In other words, the cooling slots3220are slanted at such a compound angle that the cooling slots3220become closer to the platform (16inFIG. 1A,40inFIG. 2A) at the leading edge3201than at the cooling passage3210.

Accordingly, the cooling air is exhausted through the cooling slots3220to the direction of the platform (approximately downward or centripetal direction inFIG. 6), thus film-cooling the surface of the leading edge3201of the airfoil3200.

The airfoil of gas turbine engine according to a fourth embodiment will be explained below, with reference toFIG. 7.

FIG. 7is a schematic longitudinal cross-section view of a film cooling structure at a leading edge of an airfoil of a gas turbine engine, according to a fourth embodiment.

Referring toFIG. 7, the film cooling structure at a leading edge of an airfoil of gas turbine engine according to the fourth embodiment is similar to that according to the third embodiment explained above, except for the compound angle of the cooling slots4220.

Accordingly, the compound angle of the cooling slots4220will be explained in detail below, while description of the other similar or same elements is referenced to the explanation given above.

The cooling slots4220may be slanted at such a compound angle that the cooling slots4220are farther away from the end4204of the airfoil4200at the cooling passage4210than at the leading edge4201. In other words, the cooling slots4220are slanted at such a compound angle that the cooling slots4220become farther away from the platform (16inFIG. 1A,40inFIG. 2A) at the leading edge4201than at the cooling passage4210.

Accordingly, the cooling air is exhausted through the cooling slots4220approximately in a radial direction (approximately upward direction inFIG. 7), thus film-cooling the surface of the leading edge4201of the airfoil4200.

The film cooling structure of the leading edge of the airfoil of gas turbine engine according to various embodiments provide the following advantageous effects.

According to various embodiments, since one or more cooling slots220,3200,420in elongated configuration are extended through the leading edge of the airfoil200,3200,4200to the direction of the cooling passage210,3210,4210, and across the lengthwise direction of the leading edge201,3201,4201, compared to the related art cooling holes in cylindrical configuration (38inFIG. 1A,82inFIG. 2A), the cooling efficiency of the leading edge201,3201,4201of the airfoil200,3200,4200can be greatly improved even when the location of combustion gas stagnation changes at the leading edge201,3201,4201, without having to increase the cooling air flowrate.