Gas turbine nozzles with embossments in airfoil cavities

The present application provides a nozzle for a gas turbine engine. The nozzle may include a band, a seal slot positioned within the band, an airfoil extending from the band, a cavity within the airfoil, and an embossment positioned about the band and the cavity.

TECHNICAL FIELD

The present application and the resultant patent relate generally to gas turbine engines and more particularly relate to a gas turbine nozzle with an airfoil cavity embossment so as to limit seal slot breakage into the cavity and for enhanced overall airfoil cooling.

BACKGROUND OF THE INVENTION

Impingement cooling systems have been used with turbine machinery to cool various types of components such as casings, buckets, nozzles, and the like. Impingement cooling systems cool these components via an airflow so as to maintain adequate clearances between the components and to promote adequate component lifetime. One issue with some types of known impingement cooling systems, however, is that they tend to require complicated casting and/or structural welding. Such complicated structures may not be sufficiently durable and/or they may be expensive to produce and repair.

By way of example, an impingement cooling insert may be positioned within a nozzle airfoil cavity. Nozzle ribs may be machined into the cavity for positioning the cooling insert therein. The ends of the ribs may need to be machined to achieve a proper interface for welding or brazing the cooling insert therein. Such procedures are generally time consuming and expensive. Moreover, part life may be reduced due to air leakage across the joints or otherwise.

There is thus a desire for an improved turbine nozzle. Preferably such an improved turbine nozzle may provide for the fast and efficient insertion of an impingement cooling insert therein without expensive casting or machining while providing adequate cooling for a prolonged component lifetime and overall system efficiency.

SUMMARY OF THE INVENTION

The present application and the resultant patent provide a nozzle for a gas turbine engine. The nozzle may include a band, a seal slot positioned within the band, an airfoil extending from the band, a cavity within the airfoil, and an embossment positioned about the band and the cavity.

The present application and the resultant patent further provide a method of manufacturing a nozzle for a gas turbine engine. The method may include the steps of casting an airfoil with a cavity and a band, adding an embossment with a curved configuration about the cavity and the band, and machining a seal slot into the band about the embossment.

The present application and the resultant patent further provide a nozzle for a gas turbine engine. The nozzle may include a band, a seal slot machined within the band, an airfoil extending from the band, a cavity within the airfoil, an impingement cooling insert positioned within the cavity, and an embossment cast about the band and the cavity. The embossment may have a curved configuration.

DETAILED DESCRIPTION

Referring now to the drawings, in which like numerals refer to like elements throughout the several views,FIG. 1shows a schematic view of gas turbine engine10as may be used herein. The gas turbine engine10may include a compressor15. The compressor15compresses an incoming flow of air20. The compressor15delivers the compressed flow of air20to a combustor25. The combustor25mixes the compressed flow of air20with a pressurized flow of fuel30and ignites the mixture to create a flow of combustion gases35. Although only a single combustor25is shown, the gas turbine engine10may include any number of combustors25positioned in a circumferential array or otherwise. The flow of combustion gases35is in turn delivered to a turbine40. The flow of combustion gases35drives the turbine40so as to produce mechanical work. The mechanical work produced in the turbine40drives the compressor15via a shaft45and an external load50such as an electrical generator and the like.

The gas turbine engine10may use natural gas, liquid fuels, various types of syngas, and/or other types of fuels and blends thereof. The gas turbine engine10may be any one of a number of different gas turbine engines offered by General Electric Company of Schenectady, N.Y., including, but not limited to, those such as a 7 or a 9 series heavy duty gas turbine engine and the like. The gas turbine engine10may 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.

FIG. 2shows an example of a nozzle60that may be used with the turbine40and the like. Any number of the nozzles60may be positioned in a circumferential array about a gas turbine axis. Generally described, the nozzle60may include an outer band65, an inner band70, and one or more airfoils75extending therebetween. The airfoil75may have one or more cavities80therein. A cooling insert85may be positioned in one or more of the cavities80. The cooling insert85may have a number of apertures therethrough for impingement cooling the internal wall surfaces of the airfoil75via a cooling medium such as steam and the like. The outer band65and/or the inner band70also may have one or more seal slots90formed therein. A seal may be positioned within the seal slots90of adjoining nozzles60to prevent the passage of the hot combustion gasses35therethrough. The nozzle60described herein is for the purpose of example only. Many other and different types of nozzle configurations may be used.

FIG. 3shows an example of a turbine nozzle100as may be described herein. The turbine nozzle100may include an outer band110and an inner band120(only one of which is shown). An airfoil130may extend between the outer band110and the inner band120. More than one airfoil130may be used. The airfoil130may have a leading sidewall140and a trailing sidewall150. The sidewalls140,150may define an airfoil cavity160therebetween. One or more seal slots170may be positioned in or about the outer band110and the inner band120for one or more seals to extend therebetween. The turbine nozzle100and the components thereof may have any suitable size, shape, or configuration.

The turbine nozzle100may include a cooling insert180positioned within the airfoil cavity160. The cooling insert180may include a number of apertures190therein. Any number of the apertures190may be used herein. The cooling insert180and the apertures190may have any suitable size, shape, or configuration. The apertures190of the cooling insert180allows for the passage of a cooling medium200therethrough so as to cool the sidewalls140,150and other surfaces of the airfoil130by impingement thereon. The cooling medium200may be steam and the like. Other types of cooling mediums may be used herein. Other components and other configurations may be used herein.

The turbine nozzle100also may include one or more embossments210. The embossment210may be an amount of additional airfoil material. The embossment210may be cast, extruded, or otherwise formed therein. The embossment210may be positioned about the airfoil cavity160in proximity to the seal slot170. The embossment210may have a substantially curved configuration220. The curved configuration220of the embossment210may extend into the airfoil cavity160at about a forty-five degree (45°) angle or so. Other angles and other types of curved configurations may be used herein. Specifically, the fillets and the leading angles of the embossment210may be large enough so as to enable high quality castability. Moreover, the length, the height, the size, the shape, and the configuration of the embossment210may vary. Nozzles100with varying embossments210may be used together. Other components and other configurations may be used herein.

The cooling insert180may or may not contact the embossment210. If in contact, the embossment210may serve as a last minute engagement standoff. Specifically, the standoff may be a platform for the airfoil cooling inserts. Use as the engagement standoff thus may enhance overall airfoil cooling and cooling efficiency herein.

The use of the embossment210increases the overall design space of the turbine nozzle100. Specifically, the extra material of the embossment210allows the seal slots170to be machined therein without concern of breaking into the airfoil cavity160. The seal slots170thus may have a sufficient seal depth to resist seal pull out and/or seal destruction. Breaking into the airfoil cavity160may cause a cooling flow leakage and associated reduced part life and overall performance. Moreover, a smaller overall slashface angle may be used on the inner and outer bands110,120. The use of the embossment210thus avoids a reduced seal slot depth as well as an increased slashface angle.