The present invention relates to fluid-cooled airfoils, and more particularly to fluid-cooled airfoils suitable for use with gas turbine engines.
Airfoils, such as those used in gas turbine engines, often operate in relatively hot environments. In order to help ensure air foil integrity, airfoils can utilize high temperature alloys, thermal barrier coatings, and cooling fluid delivery. However, known cooling schemes may be inadequate for some desired applications. Inadequate cooling fluid delivery can lead to spallation of coatings, and other wear or damage to the airfoil (e.g., crack formation), which may necessitate repair or replacement of the airfoil. Such a need for repair or replacement of an airfoil is costly and time-consuming. Therefore, it is desired to provide for improved fluid cooling for an airfoil, particularly at a trailing edge of the airfoil.
An apparatus according to the present invention for use with a gas turbine engine includes an airfoil, a metering opening for metering a cooling fluid, a cutback slot configured to deliver the cooling fluid from the metering opening, and a cooling hole. The airfoil defines a trailing edge, opposite first and second faces, and a mean camber line. The cutback slot is defined along the first face of the airfoil adjacent to the trailing edge and is offset from the mean camber line of the airfoil. The cooling hole has an outlet that is located at the trailing edge and substantially aligned with the mean camber line of the airfoil. The cooling hole delivers a portion of the cooling fluid from the metering opening.
In general, the present invention relates to a fluid-cooled airfoil having a film-cooling cutback slot located along a pressure face adjacent to the trailing edge and a convective-cooling hole extending to the trailing edge. A cooling fluid from a plenum is metered through a metering opening, and passes to the cutback slot to provide film cooling. A portion of the cooling fluid delivered to the cutback slot is directed through the cooling hole extending to the trailing edge to provide convective cooling to the airfoil. In that way, hybrid film cooling and convective cooling is provided at or near the trailing edge, which can help maintain regions of the trailing edge of the airfoil at or below suitable thermal operating limits. In one embodiment, an inlet of the hole extending to the trailing edge is located at or downstream from an upstream boundary of the cutback slot along the pressure face of the airfoil, and an outlet of the hole extending to the trailing edge is substantially aligned with a mean camber line of the airfoil.
Each of the cooling passages 36 (one is shown in
The trailing edge cooling hole 50 extends from the cutback slot 48 to the trailing edge 30, between an inlet 56 and an outlet 58. In the illustrated embodiment, the inlet 56 of trailing edge cooling hole 50 is located essentially within the cutback slot 48, that is, the inlet 56 is located approximately at or downstream of the upstream boundary 52 of the cutback slot 48 and at or upstream of the downstream boundary 54. Furthermore, in the illustrated embodiment, the outlet 58 of the trailing edge cooling hole 50 is substantially aligned with the mean camber line 42 at the trailing edge 30. The outlet 58 and other portions of the trailing edge cooling hole 50 has a substantially circular cross-section in the illustrated embodiment. In alternative embodiments, other shapes of the outlet 58 are possible, such as an elliptical or “racetrack” shape with a major axis arranged in the spanwise direction. The outlet 58 has a diameter (or width) D2. In one embodiment, the diameter D1 of the trailing edge 30 is at least approximately three times larger than the diameter D2 of the outlet 58. Having the diameter D1 significantly larger than the diameter D2 helps promote structural integrity of the trailing edge 30.
Although in the illustrated embodiment only a single trailing edge cooling hole 50 extends from each cutback slot 48, in further embodiments multiple trailing edge cooling holes 50 can extend from a given cutback slot 48. For example, multiple trailing edge cooling holes 50 can extend from a given cutback slot 48 at different angles relative to the centerline CL and each have separate inlets 56. Alternatively, multiple trailing edge cooling holes 50 extending from a given cutback slot 48 could share a common inlet 56.
The second portion 60B of the cooling fluid flow 60 also provides aerodynamic benefits by helping to straighten fluid flows at or near the trailing edge 30 of the airfoil 26. Moreover, by exhausting the second portion 60B of the cooling fluid flow 60 at the trailing edge 30 along the mean camber line 42, the relatively high mixing losses typically associated with pressure face and suction face cooling flows are avoided.
While the invention has been described with reference to an exemplary embodiment(s), it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the invention. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the invention without departing from the essential scope thereof. Therefore, it is intended that the invention not be limited to the particular embodiment(s) disclosed, but that the invention will include all embodiments falling within the scope of the appended claims. For example, the present invention can be utilized in conjunction with any number of additional cooling features, such as additional cooling passages of a known configuration. Moreover, trailing edge cooling holes can be drilled into existing airfoils with cutback slots as part of a repair or retrofit operation according to the present invention.
The present invention was developed, at least in part, with government funding pursuant to Contract No. N00019-02-C-3003 awarded by the United States Navy. The U.S. Government may have certain rights in this invention.
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