Patent Application
20090081029
The present invention relates in general to actively cooled devices used in high-temperature applications, and in particular to an improved cooling scheme for a gas turbine engine airfoil.
In a gas turbine engine, air is pressurized in a compressor and is mixed with fuel in a combustor for generating hot combustion gases which flow downstream through turbine stages that extract energy therefrom. The turbine includes stationary airfoils (vanes) that direct the combustion gases through respective downstream rows of rotating airfoils (blades) extending radially outwardly from a rotating shaft.
Present-day high performance turbines include vanes that are capable of withstanding temperatures approaching 1600° C. or higher. While high temperature metal alloys and ceramic materials may be used for constructing the vanes and blades, active cooling of the structures with a cooling fluid is required in many applications. Cooling is typically accomplished by directing cooling air through the hollow cavity of the airfoil.
Various schemes have been used in the past to actively cool gas turbine components such as the stationary vanes. For example, in U.S. Pat. No. 5,772,398, entitled COOLED TURBINE GUIDE VANE, a cooled turbine vane is disclosed as including a hollow aerodynamic portion between inner and outer platforms. The interior of the aerodynamic portion is partitioned into a leading edge duct and a main cavity in which a perforated tubular member is disposed, being spaced from the interior and exterior side walls of the vane by longitudinal ribs. The tubular member is divided by a partition into two cavities on the interior and exterior sides of the partition. A first cooling circuit includes the leading edge duct and the interior cavity of the tubular member, and a second cooling circuit includes the exterior cavity and a cooling system for the inner platform, both circuits being supplied with cooling air by the same source from the outer platform. Cooling air from each circuit passes through the perforations of the tubular member to impinge on the inside face of the respective side wall of the vane and is then guided toward the trailing edge, where it escapes through slits in the trailing edge wall.
Another example of a prior art device is disclosed in U.S. Pat. No. 5,813,827, entitled APPARATUS FOR COOLING A GAS TURBINE AIRFOIL. This apparatus includes two radially extending passages connected to the outer shroud to direct a cooling fluid to a plenum formed about mid-span adjacent to the trailing edge. Two arrays of cooling fluid passages extend from the plenum. One array extends radially inward toward the inner shroud. The plenum distributes the cooling fluid to the two arrays of passages so that it flows radially inward and outward to manifolds formed in the inner and outer shrouds. The manifolds direct the spent cooling fluid to a discharge passage.
To utilize the cooling air passing through a gas turbine vane effectively, it is useful to reduce the size of the cooling passage, since cooling air traveling along the center of the passage is not in contact with the surface being cooled. However, the reduction of the internal cooling passage cross-sectional area to a desired degree for cooling purposes would result in an undesirably thin aerodynamic shape or the necessity for installing complicated and costly structures within the vane for directing the fluid flow. A thinner aerodynamic shape requires a larger number of airfoils to produce the desired aero performance, thereby increasing cost and reducing efficiency. A means of improving cooling efficiency without affecting the external airfoil contour is desired.
The invention is explained in the following description in view of the drawings that show:
a is a diagram of a restraining screen for use with the structure shown in
b is a diagram of an exemplary sphere that may be used as a geometrically-shaped member in the structure shown in
a is a planar view of a restraining member for use with the structure shown in
The present inventors have discovered that a cooling flow passing through a gas turbine airfoil cooling passage can be partially blocked and caused to flow preferentially at a higher rate along the walls of the airfoil passage by filling the passage with a stacked plurality of geometric shapes. The advantageous preferential flow pattern along the walls of the passage is caused by the inability of the geometric shapes to stack as closely against the flat walls as they do stack against each other in the central portions of the passage. The stacked plurality of geometric shapes also provides a tortuous flow path for the cooling fluid, resulting in improved mixing of the fluid without the need for forming ribs or other flow disruption structures on the cooled surface. The partial blockage and preferential flow proximate the walls of the passage created by the stacked geometric shapes provide for a reduced coolant flow rate and a simultaneously improved wall cooling effect. Introducing such stacked geometric shapes into the cooling passage of a gas turbine airfoil allows the use of a sufficiently thick external contour designed for proper combustion gas path aerodynamic behavior without the usual attendant need for a very large cooling air flow or the need for expensive flow directing structures within the airfoil. The flow blockage may be formed using ceramic or metallic shapes, for example spheres, packed into the cooling cavity. The shapes may be retained in the cavity by using a perforated retaining structure such as a grate at both the inlet and outlet ends of the cavity. The retaining structure may be formed to preferentially allow more coolant to enter the cavity proximate the cooled walls, thereby further augmenting the benefits of the present invention.
Referring now to the drawings and in particular to
In accordance with the present invention, a plurality of geometrically-shaped members 27 are placed in the cavity 16, which act to direct the coolant flow 18 preferentially toward the outer walls for cooling thereof. That is, the members 27 form a partial blockage of the flow in the center of the cavity 16, which preferentially forces the coolant toward the outer walls where it is needed for cooling the hot airfoil walls. The members 27 may be any geometric shape that provides for relatively close packing against adjacent members 27 while providing relatively more open packing against an adjacent wall 144. The members may be metallic or ceramic spheres, for example. Typical metallic materials that may be used include commercially available alloys designated in the trade as IN625, IN718, Rene80, or Hastx.
With reference to
In accordance with one embodiment, the members 27 are retained within the cavity 16 by means of a screen grid 30 or other retaining structure, which may be welded to the wall 14 or otherwise supported. Details of the screen grid 30 are shown in the plan view of
The geometrically-shaped members 27 may have a constant size and shape or the size and/or shape may vary within any one cavity 16, as illustrated in
Referring now to
The embodiment illustrated in
The present invention has numerous advantages over traditional airfoil cooling schemes. First, the preferential redirection of the coolant flow to the outer surfaces of the cavities increases the efficiency of the system. Moreover, the use of simple geometrically-shaped members 27 is a cost effective means for redirecting the coolant flow without the use of elaborate and expensive duct work.
While various embodiments of the present invention have been shown and described herein, it will be obvious that such embodiments are provided by way of example only. Numerous variations, changes and substitutions may be made without departing from the invention herein. Accordingly, it is intended that the invention be limited only by the spirit and scope of the appended claims.
