Patent Application
20130011265
Not applicable.
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.
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
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.
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.
The present invention is described in detail below with reference to the attached drawing figures, wherein:
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
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
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
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
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
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
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
As it can be seen from
The gas turbine vane assembly 500 further comprises a second vane assembly 400 that is secured to the first vane assembly 200. Referring to
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
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
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
Referring to
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
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.
