US20150167475A1

US20150167475A1 - Airfoil of gas turbine engine

Info

Publication number: US20150167475A1
Application number: US14/141,943
Authority: US
Inventors: Jeong-Seek KANG; Dong-ho Rhee; Bong-Jun CHA
Original assignee: Korea Aerospace Research Institute KARI
Current assignee: Korea Aerospace Research Institute KARI
Priority date: 2013-12-17
Filing date: 2013-12-27
Publication date: 2015-06-18
Also published as: KR20150070529A; KR101565452B1

Abstract

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.

Description

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.
FIGS. 1A and 1B illustrate turbine nozzle vanes 12 disclosed in U.S. Pat. No. 7,001,141.
Referring to FIG. 1A, the related art turbine nozzle vanes 12 include an inner platform 16, an outer platform 18, and an airfoil section 20 extending radially (i.e., transversal direction in FIG. 1A) between the inner and outer platforms 16, 18 to form gas flow passage 14, for the cooling purpose. To be specific, the airfoil 20 includes film cooling holes 38 in cylindrical configuration to cool the surface of the airfoil 20, and the platforms 16, 18 have cylindrical film cooling holes 40 to cool the surfaces of the platforms.
Accordingly, as illustrated in FIG. 1A, some of the cooling air guided from the plenum regions 34, 36 cools the surfaces of the platforms 16, 28 as the air is exhausted onto the surfaces of the platforms 16, 28 through the film cooling holes 40. Referring to FIGS. 1A and 1B, the rest of the cooling air that is guided from the plenum regions 34, 36 is introduced into the airfoil cavity 24 (see FIG. 1B) and then exhausted onto the surface of the airfoil 20, thus cooling the surface of the airfoil 20. Reference numeral 22 denotes turbine blade, 26 is pressure wall of the airfoil 20, and 28 is a suction wall of the airfoil 20.
FIGS. 2A and 2B illustrate turbine blade 30 disclosed in U.S. Pat. No. 6,402,471.
Referring to FIG. 2A, the related art turbine blade 30 includes hollow platform 40, and a hollow airfoil 42 extending radially on the platform 40 (upward direction in FIG. 2A), for cooling purpose. To be specific, the airfoil 42 includes cylindrical film cooling holes to cool the surface of the airfoil 42, and the platform 40 has cylindrical film cooling holes 113 to cool the surface of the platform 40.
Accordingly referring to FIG. 2A, some of the cooling air that is guided from the hollow cavity (not illustrated) of the platform 40 is exhausted onto the surface of the platform 40 through the film cooling holes 113, thus cooling the surface of the platform 40. Referring to FIGS. 2A and 2B, the rest of the cooling air that is guided from the hollow cavity (not illustrated) of the platform 40 is introduced into the cavity (see FIG. 2B) of the airfoil 42 and then exhausted onto the surface of the airfoil 42 through the film cooling holes 82, thus cooling the surface of the airfoil 42. Reference numeral 44 denotes a pressure side of the airfoil, and 46 is 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).

BRIEF DESCRIPTION OF THE DRAWINGS

The above and/or other aspects of the present inventive concept will be more apparent by describing certain exemplary embodiments of the present inventive concept with reference to the accompanying drawings, in which:

FIG. 1 is a schematic perspective view of a related art gas turbine engine, i.e., a related art turbine nozzle vane having airfoil;

FIG. 1B is a transverse section view of the airfoil of FIG. 1A;

FIG. 2A is a schematic perspective view of a related art gas turbine engine, i.e., related art turbine blade having airfoil;

FIG. 2B is a transverse section view of the airfoil of FIG. 2A;

FIG. 3 is 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;

FIG. 4 is a cross section view of the airfoil of FIG. 3, taken on line IV-IV;

FIG. 5 is 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;

FIG. 6 is 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; and

FIG. 7 is 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.

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.
In the following description, same drawing reference numerals are used for the same elements even in different drawings. The matters defined in the description, such as detailed construction and elements, are provided to assist in a comprehensive understanding of the present inventive concept. Accordingly, it is apparent that the exemplary embodiments of the present inventive concept can be carried out without those specifically defined matters. Also, well-known functions or constructions are not described in detail since they would obscure the invention with unnecessary detail.
FIG. 3 is 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, and FIG. 4 is a cross section view of the airfoil of FIG. 3, taken on line IV-IV.
According to a first embodiment, the airfoil of gas turbine engine includes a cooling passage 210, and one or more cooling slots 220.
Referring to FIG. 4, the cooling passage 210 is formed inside the airfoil 200 to guide the cooling air from a cavity (or plenum region) (34 or 36 in FIG. 1A) of a platform (16 or 18 in FIG. 1A, 40 in FIG. 2A) into the one or more cooling slots 220. Although not illustrated, the cooling passage 210 may be divided into a plurality of spaces defined by a plurality of partitions.
Referring to FIGS. 3 and 4, the cooling slots 220 are passed through the leading edge 201 of the airfoil 200 to the direction of the cooling passage 210. To be specific, the cooling slots 220 may be elongated across the lengthwise direction (vertical direction in FIG. 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 passage 210 is passed through the leading edge 201 of the airfoil 200 to the direction of the cooling passage 210, and one or more elongated cooling slots 220 are formed across the lengthwise direction of the leading edge 201, compared to the related art cooling holes (38 in FIG. 1A, 82 in FIG. 2A) with cylindrical shape, it is possible to increase cooling efficiency of the leading edge 201 of the airfoil 200 even when the location of combustion gas stagnation varies, without having to increase the cooling air flowrate.
Further, referring to FIGS. 3 and 4, one end of the cooling slot 220 may be placed on a compressive surface 202 of the airfoil 200, while the other end thereof may be placed on a suction surface 203 of the airfoil 200. Accordingly, since the cooling slots 220 are formed across the suction surface 203, it is possible to increase cooling efficiency of the leading edge 201 of the airfoil 200 even when the location of combustion gas stagnation further changes.
Further, referring to FIG. 3, both corners 220 a of the cooling slot 220 may be rounded in the shape of circular arc. Accordingly, it is possible to prevent stress concentration from generating on both corners of the cooling slot 220.
The airfoil 220 may be the airfoil of the turbine nozzle vane (see FIG. 1A) or the airfoil of the turbine blade (see FIG. 2A).
The airfoil of the gas turbine engine according to a second embodiment will be explained below, with reference to FIG. 5.
FIG. 5 is 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 to FIG. 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 holes 230 additionally formed between the cooling slots 221, 222 of the leading edge. Accordingly, the cooling holes 230 formed between the cooling slots 221, 222 will 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 slots 220 may include first and second cooling slots 221, 222. For the explanation of the first and second cooling slots 221, 222, the explanation of the cooling slots 220 given above in the first embodiment will be referenced.
The one or more cooling holes 230 may be formed in the leading edge 201, i.e., between the first and second cooling slots 221, 222. For example, the cooling holes 230 may be formed in a known manner so that the cooling holes 230 may be cylindrical holes that are passed through to the direction of the cooling passage 210.
Since one or more elongated cooling slots 220 are provided, it is possible to increase cooling efficiency of the leading edge 201 of the airfoil 200 even when the location of combustion gas stagnation at the leading edge 201 varies, without having to increase the cooling air flowrate. Further, since one or more cylindrically-extending cooling holes 230 are provided, the same cooling function as that of the related art cooling holes (38 in FIG. 1A, 82 in FIG. 2A) may be additionally used.
The airfoil of gas turbine engine according to a third embodiment will be explained below, with reference to FIG. 6.
FIG. 6 is 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 to FIG. 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 slots 3220 that have compound angle. Accordingly, the cooling slots 3220 with 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 slots 3220 may be slanted with respect to an end 3204 of the airfoil 3200. To be specific, the cooling slots 3220 may be slanted at such a compound angle that the cooling slots 3220 become closer to the end 3204 of the airfoil 3200 at the cooling passage 3210 than at the leading edge 3201. In other words, the cooling slots 3220 are slanted at such a compound angle that the cooling slots 3220 become closer to the platform (16 in FIG. 1A, 40 in FIG. 2A) at the leading edge 3201 than at the cooling passage 3210.
Accordingly, the cooling air is exhausted through the cooling slots 3220 to the direction of the platform (approximately downward or centripetal direction in FIG. 6), thus film-cooling the surface of the leading edge 3201 of the airfoil 3200.
The airfoil of gas turbine engine according to a fourth embodiment will be explained below, with reference to FIG. 7.
FIG. 7 is 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 to FIG. 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 slots 4220.
Accordingly, the compound angle of the cooling slots 4220 will be explained in detail below, while description of the other similar or same elements is referenced to the explanation given above.
The cooling slots 4220 may be slanted at such a compound angle that the cooling slots 4220 are farther away from the end 4204 of the airfoil 4200 at the cooling passage 4210 than at the leading edge 4201. In other words, the cooling slots 4220 are slanted at such a compound angle that the cooling slots 4220 become farther away from the platform (16 in FIG. 1A, 40 in FIG. 2A) at the leading edge 4201 than at the cooling passage 4210.
Accordingly, the cooling air is exhausted through the cooling slots 4220 approximately in a radial direction (approximately upward direction in FIG. 7), thus film-cooling the surface of the leading edge 4201 of the airfoil 4200.
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 slots 220, 3200, 420 in elongated configuration are extended through the leading edge of the airfoil 200, 3200, 4200 to the direction of the cooling passage 210, 3210, 4210, and across the lengthwise direction of the leading edge 201, 3201, 4201, compared to the related art cooling holes in cylindrical configuration (38 in FIG. 1A, 82 in FIG. 2A), the cooling efficiency of the leading edge 201, 3201, 4201 of the airfoil 200, 3200, 4200 can be greatly improved even when the location of combustion gas stagnation changes at the leading edge 201, 3201, 4201, without having to increase the cooling air flowrate.
The foregoing exemplary embodiments and advantages are merely exemplary and are not to be construed as limiting the present invention. The present teaching can be readily applied to other types of apparatuses. Also, the description of the exemplary embodiments of the present inventive concept is intended to be illustrative, not to limit the scope of the claims, and many alternatives, modifications, and variations will be apparent to those skilled in the art.

Claims

What is claimed is:

1. An airfoil of gas turbine engine comprising an efficient film cooling structure at a leading edge thereof, comprising:

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.

2. The airfoil of claim 1, wherein one end of each of the one or more cooling slots is on a compressive surface of the airfoil, and the other end thereof is on a suction surface of the airfoil.

3. The airfoil of claim 1, wherein both corners of each of the one or more cooling slots are rounded to a circular arc shape.

4. The airfoil of claim 1, wherein the one or more cooling slots comprise a first cooling slot and a second cooling slot, and

the film cooling structure further comprises 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.

5. The airfoil of claim 1, wherein the one or more cooling slots are slanted with respect to an end of the airfoil.

6. The airfoil of claim 5, wherein the one or more cooling slots are so slanted that the cooling slots become closer to the end of the airfoil at the cooling passage than at the leading edge.

7. The airfoil of claim 5, wherein the one or more cooling slots are so slanted that the cooling slots become farther away from the end of the airfoil at the cooling passage than at the leading edge.

8. The airfoil of claim 1, wherein the airfoil is an airfoil of a turbine nozzle vane or a turbine blade.