Patent Application
20150075180
Embodiments of the disclosure relate generally to a gas turbine engine and more particularly relate to systems and methods for providing one or more cooling holes in a slash face of a turbine bucket.
A gas turbine engine typically includes a compressor, a combustor, and a turbine. The efficiency of the turbine depends in part on the amount of cooling air flow from the compressor that is used to cool components in the hot gas path in the turbine section. The cooling air flow may be introduced into the wheel space of the turbine to limit (or purge) high-temperature gases from entering into the wheel space. Excess purge flow to the wheel space may decrease turbine efficiency since the cooling air flow may not be available for work production.
Some or all of the above needs and/or problems may be addressed by certain embodiments of the disclosure. According to one embodiment, there is disclosed a turbine bucket. The turbine bucket may include a platform and a shank portion extending radially inward from the platform. The shank portion may include a slash face, a radial seal pin groove formed in the slash face, and at least one cooling hole disposed in the slash face about the radial seal pin groove.
According to another embodiment, there is disclosed a gas turbine engine system. The system may include a compressor, a combustor in communication with the compressor, and a turbine in communication with the combustor. The turbine bucket may include a platform and a shank portion extending radially inward from the platform. The shank portion may include a slash face, a radial seal pin groove formed in the slash face, and at least one cooling hole disposed in the slash face about the radial seal pin groove.
Further, according to another embodiment, there is disclosed a shank portion of a turbine bucket. The shank portion may include a slash face, a radial seal pin groove formed in the slash face, and at least one cooling hole disposed in the slash face about the radial seal pin groove.
Other embodiments, aspects, and features of the invention will become apparent to those skilled in the art from the following detailed description, the accompanying drawings, and the appended claims.
Reference will now be made to the accompanying drawings, which are not necessarily drawn to scale.
Illustrative embodiments will now be described more fully hereinafter with reference to the accompanying drawings, in which some, but not all embodiments are shown. The disclosure may be embodied in many different forms and should not be construed as limited to the embodiments set forth herein. Like numbers refer to like elements throughout.
The gas turbine engine 10 may use natural gas, various types of syngas, and/or other types of fuels. The gas turbine engine 10 may be anyone of a number of different gas turbine engines such as those offered by General Electric Company of Schenectady, New York and the like. The gas turbine engine 10 may have different configurations and may use other types of components. Other types of gas turbine engines also may be used herein. Multiple gas turbine engines, other types of turbines, and other types of power generation equipment also may be used herein together.
The shank portion 56 may include a slash face 66. The slash face 66 is the circumferential edge of the shank portion 58. In some instances, the leading edge of the shank portion 58 may include a forward trench cavity 68. The forward trench cavity 68 may be formed between an angle wing seal 70 and a leading edge 72 of the platform 56. The forward trench cavity 68 may provide an area where purge air from the wheelspace interfaces with the hot combustion gases. Other components and other configurations may be used herein.
In some instances, the leading edge of the shank portion 104 may include a forward trench cavity 112. The forward trench cavity 112 may be formed between an angle wing seal 114 and a leading edge 116 of the platform 102. The forward trench cavity 112 may provide an area where purge air from the wheelspace interfaces with the hot combustion gases.
In certain embodiments, the turbine bucket 102 may include a radial seal pin groove 118 formed in the slash face 110. The radial seal pin groove 118 may extend at least partially from the platform 102 to the dovetail 108. In some instances, a radial seal pin 120 (depicted in dashed lined for clarity) may be positioned within the radial seal pin groove 118. That is, each radial seal pin groove 118 may be sized and shaped to receive a radial seal pin 120 therein to facilitate sealing between adjacent shanks portions 104 when a number of turbine buckets 100 are coupled to the rotor. U.S. Patent Pub. No. 2011/0081245 and U.S. Pat. No. 7,600,972 both describe example embodiments of a radial seal pin groove and a radial seal pin and are both hereby incorporated by reference. In some instances, only the pressure side slash face and/or the suction side slash face may include the radial seal pin groove 118 and/or the radial seal pin 120. In this manner, a slash face that does not include the radial seal pin groove 118 and/or the radial seal pin 120 may still form a seal with an adjacent turbine bucket 100 that does include the radial seal pin groove 118 and/or the radial seal pin 120.
The turbine bucket 100 may include at least one cooling hole 122 disposed in the slash face 110 about the radial seal pin groove 118. The cooling hole 122 may be disposed within a pressure side slash face and/or a suction side slash face. The cooling hole 122 may be configured to provide a flow of cooling fluid (e.g., air) to the area about the radial seal pin groove 118 and/or the radial seal pin 120. For example, the cooling hole 122 may be in communication with a flow of diverted air from the compressor by way of a cooling circuit 124. Other sources of air may be used. In some instances, the cooling circuit 124 may include a number of channels 126 or the like disposed within the turbine bucket 100. In this manner, the cooling hole 122 may be in fluid communication with any one of the channels 126. The orientation, configuration, and number of cooling circuits 124 and/or channels 126 may vary.
In certain embodiments, the cooling hole 122 may be disposed in the slash face 110 about the forward trench cavity 112. That is, the cooling hole 122 may be disposed in the slash face 110 between the angle wing seal 114 and the leading edge 116 of the platform 102. Alternatively, or in addition, the cooling hole 122 may be positioned about a radial outer portion of the radial seal pin groove 118. In another instance, the cooling hole 122 may be positioned about an upstream portion of the radial seal pin groove 118 and/or a downstream portion of the radial seal pin groove 118. The cooling hole 122 may be positioned at any location about the radial seal pin groove 118. Furthermore, in some instances, the cooling hole 122 may include a number of cooling holes 122. That is, a number of cooling holes 122 may be disposed in the slash face 110 at various locations about the radial seal pin groove 118.
The location of the cooling holes 122 facilitates cooling of the area about the radial seal pin groove 118 and/or the radial seal pin 120. In turn, the forward trench cavity 112 may require less purge air, resulting in greater efficiency of the gas turbine engine.
Although embodiments have been described in language specific to structural features and/or methodological acts, it is to be understood that the disclosure is not necessarily limited to the specific features or acts described. Rather, the specific features and acts are disclosed as illustrative forms of implementing the embodiments.
