TECHNICAL FIELD

The disclosure generally relates to a cooling passage for an engine, and more specifically to a set of cooling passages for cooling a trailing edge of an airfoil.

BACKGROUND

Turbine engines, and particularly gas or combustion turbine engines, are rotary engines that extract energy from a flow of combusted gases passing through the engine and flowing over a multitude of airfoils, including stationary vanes and rotating turbine blades.

Gas turbine engines for aircraft are designed to operate at high temperatures to maximize engine efficiency, so cooling of certain engine components, such as the high pressure turbine and the low pressure turbine, can be beneficial. Typically, cooling is accomplished by ducting cooler air from the high and/or low pressure compressors to the engine components that require cooling. Temperatures in the high pressure turbine are around 1000° C. to 2000° C. and the cooling air from the compressor is around 500° C. to 700° C. While the compressor air is at a high temperature, it is cooler relative to the turbine air, and can be used to cool the turbine.

Contemporary turbine blades and other engine components generally include one or more interior cooling circuits for routing the cooling air through the engine component to cool different portions of the engine component, and can include dedicated cooling circuits for cooling different portions of the engine component.

