Patent Application
20130064680
The present invention relates generally to gas turbine engines and, more particularly, to a cooling configuration for cooling an endwall of a component, such as a vane assembly, in a gas turbine engine.
In gas turbine engines, compressed air discharged from a compressor section and fuel introduced from a source of fuel are mixed together and burned in a combustion section, creating combustion products defining a high temperature working gas. The working gas is directed through a hot gas path in a turbine section, where the working gas expands to provide rotation of a turbine rotor. The turbine rotor may be linked to an electric generator, wherein the rotation of the turbine rotor can be used to produce electricity in the generator.
In view of high pressure ratios and high engine firing temperatures implemented in modern engines, certain components, such as airfoil assemblies, e.g., stationary vane assemblies and rotating blade assemblies within the turbine section, must be cooled with cooling fluid, such as compressor discharge air, to prevent overheating of the components and to reduce thermal stress in the components.
In accordance with an aspect of the invention, a component is provided in a gas turbine engine. The component comprises an airfoil extending radially outwardly from a endwall associated with the airfoil. The endwall extends between an upstream edge and a downstream edge and defines a cool side and a gas side. A recess cavity is defined in an overhang portion extending from a location adjacent to the downstream edge toward the upstream edge. The recess cavity extends radially into the overhang portion from the cool side toward the gas side and defines a cavity surface. A plurality of grooves extend radially into the cavity surface and have an elongated dimension extending in a direction from the downstream edge toward the upstream edge.
In accordance with yet further aspects of the invention, the endwall may include opposing lateral sides extending in an axial direction between the upstream and downstream edges, and the recess cavity may extend circumferentially between the lateral sides of the endwall. Additionally, the plurality of grooves may be spaced circumferentially across the recess cavity.
The endwall may comprise a radially inner endwall and may include an inner diameter endwall post-impingement cooling chamber located adjacent to the recess cavity. A plurality of cooling passages may be provided extending from the inner diameter endwall post-impingement cooling chamber to the downstream edge. Each of the cooling passages may extend through the overhang portion and may be located between a pair of the grooves. A radially extending raised portion of the recess cavity, between each pair of grooves, may include one of the cooling passages. An inner rail may be provided extending generally circumferentially between the inner diameter endwall post-impingement cooling chamber and the recess cavity, and the cooling passages may extend through the inner rail.
The cavity surface may be located a first distance into the endwall from a peripheral radially inner surface of the endwall, and the grooves may include a groove bottom surface located a second distance radially into the endwall greater than the first distance. The cooling passages may be located radially between the groove bottom surface and the cavity surface.
The airfoil may comprise a leading edge and a trailing edge, and the trailing edge of the airfoil may be joined to the gas side of the endwall at an axial location aligned with a portion of the recess cavity.
In accordance with another aspect of the invention, a vane assembly is provided for a gas turbine engine. The vane assembly comprises an inner endwall extending between an upstream edge and a downstream edge, and defining a cool side and a gas side. An outer endwall is spaced radially outward of the inner endwall, and an airfoil extends from the inner endwall to the outer endwall and includes a leading edge and a trailing edge. An inner rail extends generally circumferentially along the inner endwall and radially inwardly of the cool side of the inner endwall. The inner endwall includes an overhang portion extending axially from a location of the inner rail. A recess cavity is defined between the inner rail and the downstream edge. The recess cavity extends radially into the overhang portion from the cool side toward the gas side and defines a cavity surface. A plurality of grooves extend radially into the cavity surface and have an elongated dimension extending in a direction from the inner rail toward the downstream edge.
Additionally, the inner endwall may include an inner diameter endwall post-impingement cooling chamber located adjacent to the recess cavity. A plurality of cooling passages may be provided extending from the inner diameter endwall post-impingement cooling chamber to the downstream edge, and the cooling passages may extend through the inner rail.
While the specification concludes with claims particularly pointing out and distinctly claiming the present invention, it is believed that the present invention will be better understood from the following description in conjunction with the accompanying Drawing Figures, in which like reference numerals identify like elements, and wherein:
In the following detailed description of the preferred embodiment, reference is made to the accompanying drawings that form a part hereof, and in which is shown by way of illustration, and not by way of limitation, a specific preferred embodiment in which the invention may be practiced. It is to be understood that other embodiments may be utilized and that changes may be made without departing from the spirit and scope of the present invention.
Referring to
Referring to
Further, it should be understood that the terms “inner”, “outer”, “radial”, “axial”, “circumferential”, and the like, as used herein, are not intended to be limiting with regard to an orientation or particular use of the elements recited for aspects of the present invention.
The airfoil structure 30 may comprise a vane assembly including first and second airfoils or vanes 32 adapted to be supported to extend radially across the flow path 28. The vanes 32 each include a generally concave sidewall 34 defining a pressure side of the vane 32, and include an opposing generally convex sidewall 36 defining a suction side of the vane 32. The sidewalls 34, 36 extend radially between an outer diameter endwall 38 and an inner diameter endwall 40, and extend generally axially in a chordal direction between a leading edge (not seen in
The inner endwall 40 includes a gas or hot side 44 facing radially outwardly toward the flow path 28, and a cool side 46 facing radially inwardly toward the center of the turbine engine 10. The hot side 44 and cool side 46 of the inner endwall 40 extend circumferentially between opposing lateral sides 48, 50 of the endwall 40, and the lateral sides 48, 50 extend axially between an upstream edge 52 and a downstream edge 54 of the inner endwall 40. As seen in
As seen in
It may be noted that, due to migration of hot gases along the vanes 32 radially inwardly from the radially outer portions toward the radially inner portions of the vanes 32, joints between the vanes 32 and the inner endwall 40 defined at fillet portions 64 adjacent to the aft portions of the vanes 32, i.e., adjacent to the trailing edges 42, experience elevated temperatures. That is, due to a trailing edge wake effect of the hot gases flowing past the vanes 32, the temperature of the aft fillet portions 64 may be substantially greater than temperatures radially outwardly from the inner endwall 40 and axially forward of the trailing edges 42. It is normally anticipated that an increased thermal stress will exist in the region where the trailing edge 42 meets the endwall 40, and it has generally been the practice to not provide the trailing edge cooling slots 62 in the areas of the trailing edges 42 closely adjacent to the endwall 40 in order to provide sufficient material to withstand the thermal stress. As a result, it has been difficult to provide effective convective cooling to the region of the junction between the trailing edge 42 and the endwall 40, i.e., at a trailing edge corner 66, which may be a further contributing factor in the formation of thermal stress at this location. In addition, different convective cooling mechanisms are provided to the endwall 40 and to the vanes 32, resulting in a differential cooling of these components which, in combination with a difference in the mass distribution of the metal forming the trailing edge corner 66 relative to the thicker or more massive endwall, may result in a substantial thermally induced strain at the trailing edge corner 66 during transient thermal cycles.
In accordance with an aspect of the invention, an overhang portion 68 of the inner endwall 40 may be configured to reduce thermal stress at the trailing edge corner 66, such as during transient thermal cycles. Referring to
The overhang portion 68 may be provided with a recess cavity 70 extending from a downstream boundary 72, adjacent to and axially spaced from the downstream edge 54, toward the upstream edge 52. The recess cavity 70 may have an upstream boundary 74 extending circumferentially and located adjacent to the inner rail 56. The recess cavity 70 may additionally extend circumferentially between lateral boundaries 76, 78 located adjacent to and circumferentially spaced from the lateral edges 48, 50. A remaining portion of the cool side 46 at the overhang portion 68 defines a peripheral radially inner surface 80 surrounding the recess cavity 70.
The recess cavity 70 is generally defined by a hollowed out area formed in the cool side 46 of the inner endwall 40, and includes a cavity surface 82 spaced radially outwardly from a plane p1 (
As seen in
Referring to
It should be noted that the ridges 92 provide sufficient material for defining the cooling passages 94, while the grooves 84 minimize or reduce an amount of material in the recess cavity 70 extending on either side of and surrounding the cooling passages 94. Removal of material of the endwall 40 to form the recess cavity 70 reduces the structural rigidity of the endwall 40, and particularly reduces the rigidity or structural stiffness of the overhang portion 68 adjacent to the trailing edge 42 of the vanes 32. Additionally, removal of the material between the cooling passages 94 to form the grooves 84 further reduces the structural rigidity of the overhang portion 68. Hence, the mass of material of the endwall 40 adjacent to the trailing edges 42 of the vanes 32 at the trailing edge corners 66 is reduced permitting a greater degree of flexure in the endwall 40, effecting a reduced material strain at this location. It should also be noted that the reduced mass of material associated with the cooling passages, i.e., in the area of the grooves 84 and ridges 92, permits greater cooling effectiveness from the cooling passages 94 which may reduce the cooling differential between the convective cooling provided in the airfoils 32 at the trailing edges 42 and the convective cooling, provided to the overhang portion 68, additionally effecting the reduced strain at and/or near the location of the trailing edge corners 66, such as may occur during thermal cycles during operation of the gas engine 10.
An inner diameter endwall post-impingement cooling chamber 96 is located adjacent to the recess cavity 70, and the inner rail 56 is located axially between the recess cavity 70 and the post-impingement cooling chamber 96, as may be seen in
The cooling passages 94 extend axially through or past the inner rail 56 to the post-impingement cooling chamber 96. The cooling air metering through the impingement cooling holes 100, comprising post-impingement air, may pass into entry openings 106 of the cooling passages 94 and flow through the cooling passages 94 to exit openings 108 (
It should be understood that, although the recess cavity 70 illustrating aspects of the present invention is shown as a rectangular cavity extending substantially the axial and circumferential extent of the overhang portion 68, other configurations of the recess cavity 70 may be provided. For example, the overhang portion 68 may be provided with one or more recess cavities configured or shaped to address particular structural rigidity and cooling requirements associated with a specific airfoil structure 30.
While particular embodiments of the present invention have been illustrated and described, it would be obvious to those skilled in the art that various other changes and modifications can be made without departing from the spirit and scope of the invention. It is therefore intended to cover in the appended claims all such changes and modifications that are within the scope of this invention.
