Patent Application
20060013685
The present invention relates generally to gas turbine engines and more specifically to a turbine vane configuration having reduced airfoil stresses.
A gas turbine engine typically comprises a multi-stage compressor, which compresses air drawn into the engine to a higher pressure and temperature. A majority of this air passes to the combustors, which mix the compressed heated air with fuel and contain the resulting reaction that generates the hot combustion gases. These gases then pass through a multi-stage turbine, which, in turn drives the compressor, before exiting the engine. A portion of the compressed air from the compressor bypasses the combustors and is used to cool the turbine blades and vanes that are continuously exposed to the hot gases of the combustors. In land-based gas turbines, the turbine is also coupled to a generator for generating electricity.
Turbines are typically comprised of alternating rows of rotating and stationary airfoils. The stationary airfoils, or vanes, direct the flow of hot combustion gases onto the subsequent row of rotating airfoils, or blades, at the proper orientation such as to maximize the output of the turbine. As a result of the hot combustion gases passing through the vanes, the vanes operate at a very high temperature, typically beyond the capability of the material from which they are made. In order to lower the operating temperatures of the vane material to a more acceptable level, vanes are often cooled, either by air or steam. Typically, turbine vanes are configured in multiple segments, with each segment including a plurality of vanes. This configuration is well known in order to minimize hot gas leakage between adjacent vanes, thereby lowering turbine performance. While this configuration is advantageous from a leakage perspective, it has inherent disadvantages as well, including an increased stiffness along the platform that connects the adjacent vanes, relative to a single vane configuration.
A vane assembly 10 of the prior art, is shown in
What is needed is a vane assembly configuration that lowers the operating stresses in the vane and inner platform for a vane assembly having an inner rail portion that is exposed to lower operating temperatures than the platform or vane.
A turbine vane assembly is disclosed having lower thermally induced stresses in the airfoil and platform region resulting in improved component durability. The vane assembly comprises a first platform, second platform in spaced relation to the first platform, and at least one airfoil extending therebetween. The source of cracking in prior art vane assemblies related to the significant temperature differences over a short distance between the vane, platform, and inner rail, located along the inner platform, opposite to the airfoil. In the present invention, the first platform further comprises an inner rail having a rail length, a rail height, a rail thickness, an inner rail wall, and at least one opening extending from the inner rail wall and through the rail thickness. The at least one opening is sized to allow the inner platform to have reduced resistance to thermal deflections while not compromising the structural integrity of the inner platform nor allowing leakage of vane cooling fluid. Multiple embodiments of opening geometry are disclosed depending on stress reduction requirements and platform/inner rail geometry.
It is an object of the present invention to provide a turbine vane assembly having reduced thermal stresses in the airfoil and platform regions.
It is another object of the present invention to provide a turbine vane assembly having increased flexibility along the inner platform region.
In accordance with these and other objects, which will become apparent hereinafter, the instant invention will now be described with particular reference to the accompanying drawings.
The present invention is shown in detail in
The focus of the present invention is directed towards the inner rail and at least one opening located therein, such that the stress relief provided to inner rail 23 by opening 28 could be applied to a variety of vane assemblies and is not limited to the embodiment disclosed. Opening 28 is configured to allow inner platform 21 to have increased flexibility while not compromising the structural integrity of inner platform 21. For example, in the preferred embodiment of the present invention, opening 28 comprises a slot having a generally circular end, as shown in
An additional feature of the present invention is a removable seal 31 that is placed within the slot of opening 28 in order to seal inner rail 23 from any leakages of cooling fluid that is dedicated for airfoils 30. Seal 31 is fixed to inner rail 23 by a removable means such as tack welding at one end of the seal, such that the structural freedom intended by opening 28 is maintained. Seal 31, as shown in
While the invention has been described in what is known as presently the preferred embodiment, it is to be understood that the invention is not to be limited to the disclosed embodiment but, on the contrary, is intended to cover various modifications and equivalent arrangements within the scope of the following claims.
