A gas turbine engine generally includes a fan and a core arranged in flow communication with one another. Additionally, the core of the gas turbine engine generally includes, in serial flow order, a compressor section, a combustion section, a turbine section, and an exhaust section. In operation, air is provided from the fan to an inlet of the compressor section where one or more axial compressors progressively compress the air until it reaches the combustion section. Fuel is mixed with the compressed air and burned within the combustion section to provide combustion gases. The combustion gases are routed from the combustion section to the turbine section. The flow of combustion gases through the turbine section drives the turbine section and is then routed through the exhaust section, e.g., to atmosphere.
In general, turbine performance and efficiency may be improved by increased combustion gas temperatures. However, increased combustion temperatures can negatively impact the gas turbine engine components, for example, by increasing the likelihood of material failures. Thus, while increased combustion temperatures can be beneficial to turbine performance, some components of the gas turbine engine may require cooling features or reduced exposure to the combustion gases to decrease the negative impacts of the increased temperatures on the components.
Film cooling gas turbine engine components, e.g., by directing a flow of cooler fluid over the surface of the component, can help reduce the negative impacts of elevated combustion temperatures. For example, cooling apertures may be provided throughout a component that allow a flow of cooling fluid from within the component to be directed over the outer surface of the component. However, multiple rows of cooling holes often are required to achieve beneficial film cooling, and the multiple rows of cooling holes can be detrimental to the component structure as well as engine performance. Also, typical drilling processes for defining the cooling holes require increased component thicknesses to accommodate tolerances in drill hole placement, thereby increasing the weight of and material required to produce the component. Further, known cooling hole configurations often have only a single solution for metering the flow of cooling fluid.
Therefore, improved cooling features for gas turbine components that overcome one or more disadvantages of existing cooling features would be desirable. In particular, an airfoil for a gas turbine engine having trailing edge cooling features that minimize a thickness of a trailing edge portion of the airfoil would be beneficial. Moreover, an airfoil for a gas turbine engine having trailing edge cooling features that reduce cooling flow would be desirable. Further, an airfoil having trailing edge cooling features that minimize or reduce manufacturing time and cost would be advantageous. Additionally, a method for forming an airfoil for a gas turbine engine where the airfoil has features for improved trailing edge cooling would be useful.
USPTO Backgrounds
USPTO Backgrounds