US20130011265A1

US20130011265A1 - Chevron platform turbine vane

Info

Publication number: US20130011265A1
Application number: US13/176,200
Authority: US
Inventors: Edward Len Miller; Yong Wang; Stephen Wayne Fiebgier
Original assignee: Alstom Technology AG
Current assignee: General Electric Technology GmbH
Priority date: 2011-07-05
Filing date: 2011-07-05
Publication date: 2013-01-10

Abstract

Gas turbine vanes of a gas turbine vane assembly have a reconfigured platform shape so as to reduce the ingestion of hot combustion gases into platform gaps between adjacent vane assemblies is disclosed. The vane platforms are configured so the joint between adjacent vanes is repositioned a sufficient distance away from the leading edge of the vane airfoil and the associated bow wave generated by the hot combustion gases passing over the airfoil leading edge.

Description

CROSS-REFERENCE TO RELATED APPLICATIONS

Not applicable.

TECHNICAL FIELD

The present invention relates to gas turbine engines. More particularly, embodiments of the present invention relate to a gas turbine vane having a platform shaped in order to reduce ingestion of hot combustion gases into joints between adjacent vanes of a vane assembly.

BACKGROUND OF THE INVENTION

A gas turbine engine operates to produce mechanical work or thrust. For a land-based gas turbine engine, a generator is typically coupled to the engine through an axial shaft, such that the mechanical work of the engine is harnessed to generate electricity. A typical gas turbine engine comprises a compressor, at least one combustor, and a turbine, with the compressor and turbine coupled together through the axial shaft. In operation, as air passes through multiple stages of axially-spaced rotating blades and stationary vanes of the compressor, its pressure increases. The compressed air is then mixed with fuel in the combustion section, which can comprise one or more combustion chambers. The fuel-air mixture is ignited in the combustion chamber(s), producing hot combustion gases, which pass into the turbine causing the turbine to rotate. The turning of the shaft also drives the generator.
A prior art turbine vane 100 is shown in FIG. 1. The turbine vane 100 includes an inner platform 102, an outer platform 104 spaced a distance radially outward relative to an engine centerline. Positioned between and connected to the platforms 102 and 104 is at least one airfoil 106. In operation hot combustion gases pass through the channels created between the airfoils 106.
The turbine comprises a plurality of rotating and stationary stages of airfoils. For the turbine vanes, the leading edge region of the airfoil and vane platform is subjected to the aerodynamic loads from the preceding stage of turbine blades or the exit flow of a combustor. The combustion gases then pass around the airfoil, beginning at the airfoil's leading edge. Depending on the shape of the airfoil and the angle at which the flow of hot gases are imparted onto the leading edge of the airfoil, a bow wave can be created, which is an area of high pressure combustion gases extending a distance away from the airfoil leading edge. This wave of combustion gases is often forced into the region between adjacent turbine vanes in a vane assembly. Depending on the supply pressure of the cooling air within the platform region and the strength of the bow wave, the hot combustion gases of the bow wave may penetrate into the joint between adjacent vanes, causing overheating and erosion of the platform.

SUMMARY

Embodiments of the present invention are directed towards gas turbine vanes and a gas turbine vane assembly. In an embodiment of the present invention, a gas turbine vane comprises an inner arc-shaped platform, an outer arc-shaped platform, and an airfoil extending therebetween. The inner arc-shaped platform has a pressure side radial face and a suction side radial face where the pressure side radial face is formed in two intersecting portions and includes a relief cut at the intersection of the two portions. The outer arc-shaped platform, which is spaced a radial distance from the inner platform also has a pressure side radial face and suction side radial face where the pressure side radial face is also formed having two intersecting portions and includes a relief cut at the intersection of the two portions. The outer arc-shaped platform is separated from the inner arc-shaped platform by at least one airfoil. The suction side radial faces each have a generally planar wall.
In an alternate embodiment of the present invention, a gas turbine vane comprises an inner arc-shaped platform, an outer arc-shaped platform, and an airfoil extending between. The inner arc-shaped platform has a pressure side radial face and a suction side radial face where the suction side radial face is formed in two intersecting portions. The outer arc-shaped platform, which is spaced a radial distance from the inner platform also has a pressure side radial face and suction side radial face where the suction side radial face is also formed in two intersecting portions. The outer arc-shaped platform is spaced radially from the inner arc-shaped platform by at least one airfoil.
In yet another alternate embodiment of the present invention a gas turbine vane assembly is disclosed comprising a first vane assembly, a second vane assembly, and a fastener mechanism. The first vane assembly has a first inner platform with a pressure side radial face having a first portion and a second portion, a first outer platform with a pressure side radial face also having a first portion and second portion, and a first airfoil extending between the first inner platform and first outer platform. The second vane assembly has a second inner platform with a suction side radial face having a first portion and a second portion, a second outer platform with a suction side radial face also having a first portion and second portion. The first vane assembly and second vane assembly are fastened together along the surfaces opposite of the multi-surface platform faces by a fastener mechanism.
Additional advantages and features of the present invention will be set forth in part in a description which follows, and in part will become apparent to those skilled in the art upon examination of the following, or may be learned from practice of the invention.

BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWINGS

The present invention is described in detail below with reference to the attached drawing figures, wherein:

FIG. 1 depicts a perspective view of a gas turbine vane assembly of the prior art;

FIG. 2 depicts a perspective view of a gas turbine vane assembly in accordance with an embodiment of the present invention;

FIG. 3 depicts a perspective view of the inner platform of the gas turbine vane assembly of FIG. 2 in accordance with an embodiment of the present invention;

FIG. 4 depicts a perspective view of components forming a gas turbine vane assembly in accordance with an embodiment of the present invention;

FIG. 5 depicts a perspective view of the outer platforms of a gas turbine vane assembly in accordance with an embodiment of the present invention;

FIG. 6 depicts a top view of a gas turbine vane assembly in accordance with an embodiment of the present invention;

FIG. 7 depicts a graph of the gap pressure between the inner platforms of vanes of a gas turbine vane assembly of the prior art and an embodiment of the present invention;

FIG. 8 depicts a graph of the gap pressure between the outer platforms of vanes of a gas turbine vane assembly of the prior art and an embodiment of the present invention; and

FIG. 9 depicts a cross section view through a portion of the turbine vane assembly of the prior art within a turbine section showing the pressure isolines across the axial length of the turbine vane assembly.

DETAILED DESCRIPTION

The subject matter of the present invention is described with specificity herein to meet statutory requirements. However, the description itself is not intended to limit the scope of this patent. Rather, the inventors have contemplated that the claimed subject matter might also be embodied in other ways, to include different components, combinations of components, steps, or combinations of steps similar to the ones described in this document, in conjunction with other present or future technologies.
A gas turbine vane assembly 200 in accordance with an embodiment of the present invention is depicted in FIGS. 2-6. Referring to FIG. 2, the gas turbine vane assembly 200 comprises an inner arc-shaped platform 202 having a pressure side radial face 204 and a suction side radial face 206. The pressure side radial face comprises a first portion 204A, a second portion 204B, and a relief cut 208 located at the intersection of the first portion 204A and the second portion 204B.
The gas turbine vane assembly 200 further comprises an outer arc-shaped platform 210 spaced a distance radially outward of the inner arc-shaped platform 202. The outer arc-shaped platform 210 has a pressure side radial face 212 and a suction side radial face 214. The pressure side radial face 212 comprises a first portion 212A, a second portion 212B and a relief cut 216 at the intersection of the first portion 212A and second portion 212B. The gas turbine vane assembly 200 also comprises at least one airfoil 218 extending between the inner arc-shaped platform 202 and the outer arc-shaped platform 210. Although a variety of manufacturing techniques can be used, for ease of manufacturing and structural integrity, it is preferred that the inner arc-shaped platform, airfoil, and outer arc-shaped platform are integrally cast together.
The first portion 204A of the inner arc-shaped platform 202 is generally co-planar with the first portion 212A of the outer arc-shaped platform 210. Further, the second portion 204B of the inner arc-shaped platform 202 is also generally co-planar with the second portion 212B of the outer arc-shaped platform 210. Alignment of these surfaces is necessary to aid in assembly of the gas turbine vane assembly 200, as discussed below.
When the gas turbine vanes are assembled together in the turbine along their corresponding chevron portions, it is necessary to place one or more seals between adjacent platforms of the turbine vanes in order to prevent leakage between adjacent vanes. Referring again to FIG. 2, in an embodiment of the present invention, the first portion 204A and the second portion 204B of the inner arc-shaped platform 204 further comprises an inner seal slot 220. In this embodiment, the first portion 212A and second portion 212B of the outer platform 210 also comprises an outer seal slot 222. Therefore, one or more seals (not shown) can be placed in the slots 220 and 222 to seal the pressure side of the inner and outer platforms against an adjacent turbine vane. A variety of seal materials can be used, but one such material is a sheet metal seal, such as that disclosed in U.S. Pat. No. 7,334,800.
Because of the extreme operating temperatures to which the turbine vane 200 is exposed, it is often necessary to provide additional measures to help protect the turbine vane. Therefore, an embodiment of the invention includes applying a thermal barrier coating to the gas path surfaces of the inner arc-shaped platform, the outer arc-shaped platform, and at least one airfoil. Also, in an embodiment of the invention, the vane assembly may be actively cooled by directing an air source to the airfoil 218 through the outer arc-shaped platform 210.
Referring to FIGS. 2, 3, and 5, an alternate embodiment of the present invention is depicted. In the alternate embodiment, a gas turbine vane 400 comprises an inner arc-shaped platform 402 having a pressure side radial face 404 and a suction side radial face 406, where the suction side radial face has a first portion 406A and a second portion 406B. The gas turbine vane 400 also comprises an outer arc-shaped platform 408 spaced a distance radially outward of the inner arc-shaped platform 402. The outer arc-shaped platform 408 has a pressure side radial face 410 and a suction side radial face 412 where the suction side radial face 412 comprises a first portion 412A and a second portion 412B. At least one airfoil 414 extends between the inner arc-shaped platform 402 and the outer arc-shaped platform 408. The inner arc-shaped platform 402, outer arc-shaped platform 408, and airfoil 414 are preferably integrally cast together.
The first portion 406A of the inner arc-shaped platform 402 is generally co-planar with the first portion 412A of the outer arc-shaped platform 408. While the second portion 406B of the inner arc-shaped platform 402 is generally parallel to the second portion 412B of the outer arc-shaped platform 408. Furthermore, the first portion 406A and the second portion 406B of the inner arc-shaped platform 402 further comprises an inner seal slot 416. Also, the first portion 412A and second portion 412B of the outer arc-shaped platform 412 further comprises an outer seal slot 418. Similar to the first vane assembly 200, one or more sheet metal seals can be placed in slots 416 and 418 to reduce leakage along the platform sidefaces between the first and second portions of adjacent radial faces.
The gas turbine vane 400 also includes one or more alternatives for improving the thermal capability of the vane. One such alternative is a bond coating and thermal barrier coating. The bond coating and thermal barrier coating is applied to a portion of the inner arc-shaped platform, a portion of the outer arc-shaped platform and the at least one airfoil extending between the platforms. An additional way of improving thermal capability is through active cooling. The gas turbine vane 400 also comprises an airfoil 414 that is air cooled by a source of air entering the airfoil 414 through the outer arc-shaped platform 408 and passing along the walls of the airfoil and then through a plurality of openings 420 (see FIGS. 3-5).
A common prior art vane assembly configuration includes two parallel mate face surfaces, often times cut along an angle relative to the vane platform leading face, as depicted by A in FIG. 9. Depending on manufacturing and assembly tolerances, a small gap may be present between the adjacent platforms, thereby providing a way for hot combustion gases to enter the gap region. When this occurs, the hot gases, often upwards of approximately 2000 deg. F. can cause overheating and erosion to the platform leading edge region. One instance where this is known to occur, as shown in FIG. 9, is when there is a bow wave BW of hot gases extending away from the airfoil leading edge region 218A and the airfoil leading edge 218A is close enough to the vane platform edge A that because of the bow wave BW high pressure, it can enter the gap between adjacent platforms. As shown in FIG. 9, in the vane assembly the BW is shown coming off the leading edge region of each airfoil.
In yet another embodiment of the present invention, a gas turbine vane assembly is disclosed. The gas turbine vane assembly 500 is shown in detail in FIGS. 2-6. The assembly 500 comprises a first vane assembly 200, a second vane assembly 400, and a fastener mechanism 600.
Through an embodiment of the present invention, where the first vane assembly 200 is secured to the second vane assembly 400 so as to form the gas turbine vane assembly 500, significant improvements in eliminating injection of the bow wave gases is achieved, resulting in extended component life of the vane assembly 500. Referring back to FIG. 2, the gas turbine vane assembly 500 comprises a first vane assembly 200 having a first inner platform 202 having a pressure side radial face 204 with first portion 204A, second portion 204B, and a relief cut 208 at their intersection. Inner platform 202 also includes a suction side radial face 206, which is a generally straight surface. The first vane assembly 200 also comprises a first outer platform 210 that is spaced a radial distance from the first inner platform 202 and having a pressure side radial face 212 and a suction side radial face 214, where the pressure side radial face 212 has a first portion 212A and a second portion 212B joined together at their intersection by a relief cut 216. A first airfoil 218 extends between and joins together the first inner platform 202 and first outer platform 210.
As it can be seen from FIGS. 2, 3, and 5, referring to the first vane assembly 200, the first portion 204A of the inner arc-shaped platform 202 is generally co-planar with the first portion 212A of the outer arc-shaped platform 210. Furthermore, the second portion 204B of the inner arc-shaped platform 202 is also generally co-planar with the second portion 212B of the outer arc-shaped platform 210. Referring to FIGS. 3 and 5, a similar co-planar orientation of inner and outer platform mate face surfaces also exists for the suction side of the vane assemblies.
The gas turbine vane assembly 500 further comprises a second vane assembly 400 that is secured to the first vane assembly 200. Referring to FIG. 3, the second vane assembly 400 comprises a second inner platform 402 having a pressure side radial face 404 and a suction side radial face 406, where the suction side radial face 406 has a first portion 406A and a second portion 406B. The second vane assembly 400 also comprises a second outer platform 408, which is spaced a radial distance from the second inner platform 402 and also has a pressure side radial face 410 and a suction side radial face 412, where the suction side radial face 412 has a first portion 412A and a second portion 412B. The second vane assembly 400 also includes a second airfoil 414 extending between the second inner platform 402 and the second outer platform 408.
The first vane assembly 200 and second vane assembly 400 is secured together by a fastener mechanism proximate the inner and outer platforms, as shown in FIGS. 2, 4, and 5. An outer fastening mechanism 600 includes a first bracket 602 secured to a first vane assembly 200 and a second bracket 604 secured to a second vane assembly 400. The first bracket 602 and second bracket 604 are held together by a removable fastener such as a pin or bolt (not shown). Referring to FIG. 3, an inner fastening mechanism 606 is shown and includes brackets 608 and 610 Like the outer fastening mechanism 600, the inner fastening mechanism 606 is also utilizes removable fasteners, such as a pin or bolt (also not shown) for securing the brackets 608 and 610 together. Because of the vane assembly orientation, the outer fastening mechanism 600 utilizes three fasteners, where the inner fastening mechanism 606 utilizes one fastener.
The vane assembly 500 is oriented such that the suction side radial face 214 of the outer platform 210 of the first vane assembly 200 is adjacent to the pressure side radial face 410 of the outer platform 408 of second vane assembly 400, as shown in FIG. 2. The fastener mechanism 600 secures the first vane assembly 200 to the second vane assembly 400 so as to form a sealed interface. Because of the alternate platform configuration disclosed above, and the fastening mechanisms 600 and 606, the joint between the platforms is better sealed than the prior art configurations and where the joint is not bolted together. For the non-bolted mateface, a gap exists and seals are used. With the chevron configuration disclosed herein, the gap has been moved further away from any bow wave coming off the airfoil leading edge, such that the hot combustion gases do not enter the gap between the vane assemblies.
While the embodiments of the present invention improve the sealing between adjacent vanes of a vane assembly, any hot combustion gases that do leak between the platform surfaces can be minimized through alternate sealing arrangements. A plurality of flexible sheet metal seals (not shown) can be positioned in slots of the inner and outer platforms to prevent the flow of gases or compressed air in between any gaps of the platforms. More specifically, and as shown in FIGS. 2 and 3, the first vane assembly 200, includes a series of slots in the platform mate faces, such as slots 220 in the first inner platform 202 and slots 222 in the outer platform 210.
Referring to FIG. 7, a chart showing static gap pressure along the mateface of the inner diameter platforms is shown. The solid line (without symbols) indicates the pressure of the cooling air under the platform. The line with the solid data points shows the gap pressure of the prior art vane of FIG. 1. As it can be seen from FIG. 7, the prior art vane assembly has a higher gap pressure than the underplatform pressure along part of the platform. This indicates that some of the hot combustion gases enters the platform gap (as caused by the bow wave and shape of vane platform), thereby causing erosion in the platform. The shaded data points indicate the gap pressure of the vane assembly 500 of the present invention. As it can be seen from FIG. 7, the inner vane platform gap pressure is lower than the under platform pressure. In this embodiment, because the inner vane platform gap pressure is lower than the under platform pressure, hot combustion gases do not enter the gap between adjacent platforms.
While the present invention corrects an inflow problem along the inner arc-shaped platforms, no significant inflow problem exists at the outer platform for the prior art configuration, as shown in FIG. 8. However, the outer platform design of the present invention further increases the margin between the underplatform pressure side and the static pressure at the platform gap.
The present invention has been described in relation to particular embodiments, which are intended in all respects to be illustrative rather than restrictive. Alternative embodiments will become apparent to those of ordinary skill in the art to which the present invention pertains without departing from its scope.
From the foregoing, it will be seen that this invention is one well adapted to attain all the ends and objects set forth above, together with other advantages which are obvious and inherent to the system and method. It will be understood that certain features and sub-combinations are of utility and may be employed without reference to other features and sub-combinations. This is contemplated by and within the scope of the claims.

Claims

1. A gas turbine vane comprising:

an inner arc-shaped platform having a pressure side radial face and a suction side radial face where the pressure side radial face comprises a first portion, a second portion, and a relief cut at the intersection of the first portion and the second portion;

an outer arc-shaped platform spaced a distance radially outward of the inner arc-shaped platform and having a pressure side radial face and a suction side radial face where the pressure side radial face comprises a first portion, a second portion, and a relief cut at the intersection of the first portion and the second portion; and,

at least one airfoil extending between the inner arc-shaped platform and the outer arc-shaped platform.

2. The gas turbine vane of claim 1, wherein the first portion of the inner arc-shaped platform is generally co-planar with the first portion of the outer arc-shaped platform and the second portion of the inner arc-shaped platform is generally co-planar with the second portion of the outer arc-shaped platform.

3. The gas turbine vane of claim 1, wherein the first portion and the second portion of the pressure side radial face of the inner arc-shaped platform further comprises an inner seal slot.

4. The gas turbine vane of claim 1, wherein the first portion and the second portion of the pressure side radial face of the outer arc-shaped platform further comprises an outer seal slot.

5. The gas turbine vane of claim 1 further comprising a bond coating and a thermal barrier coating applied to a portion of the inner arc-shaped platform, a portion of the outer arc-shaped platform, and the at least one airfoil.

6. The gas turbine vane of claim 1, wherein the airfoil is air cooled by a source of air entering the airfoil through the outer arc-shaped platform.

7. The gas turbine vane of claim 1, wherein the inner arc-shaped platform, the outer arc-shaped platform, and the airfoil are integrally cast.

8. A gas turbine vane comprising:

an inner arc-shaped platform having a pressure side radial face and a suction side radial face where the suction side radial face comprises a first portion and a second portion;

an outer arc-shaped platform spaced a distance radially outward of the inner arc-shaped platform and having a pressure side radial face and a suction side radial face where the suction side radial face comprises a first portion and a second portion; and,

at least one airfoil extending between the inner arc-shaped platform and the outer arc-shaped platform;

9. The gas turbine vane of claim 8, wherein the first portion of the inner arc-shaped platform is generally co-planar with the first portion of the outer arc-shaped platform and the second portion of the inner arc-shaped platform is generally co-planar with the second portion of the outer arc-shaped platform.

10. The gas turbine vane of claim 8, wherein the first portion and the second portion of the suction side radial face of the inner arc-shaped platform further comprise a seal slot.

11. The gas turbine vane of claim 8, wherein the first portion and the second portion of the suction side radial face of the outer arc-shaped platform further comprise a seal slot.

12. The gas turbine vane of claim 8 further comprising a bond coating and a thermal barrier coating applied to a portion of the inner arc-shaped platform, a portion of the outer arc-shaped platform, and the at least one airfoil.

13. The gas turbine vane of claim 8, wherein the airfoil is air cooled by a source of air entering the airfoil through the outer arc-shaped platform.

14. The gas turbine vane of claim 8, wherein the inner arc-shaped platform, the outer arc-shaped platform, and the airfoil are integrally cast.

15. A gas turbine vane assembly comprising:

a first vane assembly having:

a first inner platform having a pressure side radial face and a suction side radial face, where the pressure side radial face has a first portion, a second portion, and a relief cut at the intersection of the first and second portion;

a first outer platform spaced a radial distance from the first inner platform and having a pressure side radial face and a suction side radial face, where the pressure side radial face has a first portion and a second portion, and a relief cut at the intersection of the first and second portion; and,

a first airfoil extending between the first inner platform and first outer platform;

a second vane assembly having:

a second inner platform having a pressure side radial face and a suction side radial face, where the suction side radial face has a first portion and a second portion;

a second outer platform spaced a radial distance from the second inner platform and having a pressure side radial face and a suction side radial face, where the suction side radial face has a first portion and a second portion; and,

a second airfoil extending between the second inner platform and second outer platform; and,

a fastener mechanism positioned adjacent the first and second outer platforms and the first and second inner platforms for securing the first vane assembly to the second vane assembly.

16. The gas turbine vane assembly of claim 15, wherein the fastener mechanism comprises a first bracket secured to the first vane assembly and a second bracket secured to the second vane assembly and one or more removable fasteners.

17. The gas turbine vane assembly of claim 16, wherein the suction side radial face of the outer platform of the first vane assembly is secured to the pressure side radial face of the outer platform of the second vane assembly by the fastener mechanism.

18. The gas turbine vane assembly of claim 17, wherein the fastener mechanism causes the first vane assembly to contact the second vane assembly so as to form a sealed interface.

19. The gas turbine vane assembly of claim 17, wherein the first portion of the inner arc-shaped platform is generally co-planar with the first portion of the outer arc-shaped platform and the second portion of the inner arc-shaped platform is generally co-planar with the second portion of the outer arc-shaped platform.

20. The gas turbine vane assembly of claim 15 further comprising a plurality of flexible sheet metal seals positioned in respective slots of the inner and outer platforms to provide a seal between adjacent turbine vane assemblies.