Turbine blade trailing edge with low flow framing channel

#The present invention relates to a cooling system for use in an airfoil of a turbine engine, and more particularly, to a trailing edge cooling circuit and core used for forming the same.

#In a gas turbine engine, compressed air discharged from a compressor section is mixed with fuel and burned in a combustion section, creating combustion products comprising hot combustion gases. The combustion gases are directed through a hot gas path in a turbine section comprising a series of turbine stages typically including a plurality of paired rows of stationary vanes and rotating turbine blades. The turbine blades extract energy from the combustion gases and provide rotation of a turbine rotor for powering the compressor and providing output power.

#The airfoils of the vanes and blades are typically exposed to high operating temperatures, and thus include cooling circuits to remove heat from the airfoil and to prolong the life of the vane and blade components. A portion of the compressed air discharged from the compressor section may be diverted to these cooling circuits. Manufacture of airfoils with one or more cooling circuits typically requires the use of a ceramic core comprising framing channels at the radially inner and outer portions in order to provide sufficient structural stability and to prevent unzipping of the ceramic core during casting.

#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:

#FIG. 1 is a perspective view of an airfoil assembly according to the present invention in which a portion of the outer wall is cut away to illustrate aspects of the invention in detail;

#FIGS. 2A and 2B are enlarged side views of the sections indicated by boxes 2A and 2B, respectively, in FIG. 1;

#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.

#The present invention provides a construction for an airfoil located within a turbine section of a gas turbine engine (not shown). Referring now to FIG. 1, an exemplary airfoil assembly 10 constructed in accordance with an aspect of the present invention is illustrated. The airfoil assembly 10 includes an airfoil 11, a platform 17, and a root 18 that is used to conventionally secure the airfoil assembly 10 to a shaft and disc assembly of the turbine section (not shown) for supporting the airfoil assembly 10 in the gas flow path of the turbine section. Although aspects of the invention are discussed herein with specific reference to components of a blade assembly in a gas turbine engine, those skilled in the art will understand that the concepts disclosed herein could also be used in the formation of a stationary vane assembly.

#The airfoil 11 shown in FIG. 1 includes an outer wall defining a leading edge 12, a trailing edge 13, a suction side 20, a pressure side (not labeled) opposite the suction side 20, a radially inner end 15 adjacent to the platform 17, and a radially outer tip 22. As used throughout, unless otherwise noted, the terms “radial,” “radially inner,” “radially outer,” and derivatives thereof are used with reference to a radial direction as represented by arrow R in FIG. 1, which is parallel to a longitudinal axis of the airfoil 11. The terms “axial,” “upstream,” “downstream,” and derivatives thereof are used with reference to a flow of combustion gases through the hot gas path in the turbine section, and an “axial direction” is defined between the leading and trailing edges 12, 13 of the airfoil 11. The airfoil 11 extends in a radial direction R from the radially inner end 15 to the radially outer tip 22.