Abstract: A cooled airfoil includes an impingement rib having a multiple of openings which supply a cooling airflow from a cooling circuit flow path toward an airfoil leading edge. The multiple of openings are offset in the impingement rib opposite an outer airfoil wall which includes gill holes. Offsetting the multiple of openings opposite an outer airfoil wall which includes the gill holes focuses the cooling airflow across turbulators to increase the cooling airflow dwell time to increase the thermal transfer therefrom in higher temperature airfoil areas.

Abstract: An airfoil assembly includes an airfoil with at least one cavity that is in communication with a source of cooling air. A baffle is disposed within that cavity and includes a plurality of openings for directing cooling air against the hot wall. A plurality of dividers extends between the baffle walls and the internal cavity to direct cooling air toward one of a leading edge chamber and a trailing edge chamber.

Abstract: A vane for use in a gas turbine engine has a platform connected to an airfoil. There is a cooling passage for supplying cooling air to the platform. A cooling chamber supplies cooling air to a plurality of cooling slots at the platform. The cooling slots have a non-uniform cross section.

Abstract: A vane cluster has a coated metallic substrate. The cluster includes a platform and a shroud. At least first and second airfoils extend between an outer face of the platform and an inner face of the shroud. Each airfoil has a pressure side and a suction side. The pressure side of the first airfoil faces the suction side of the second airfoil. The cluster includes a cooling passageway system including one or more first feed passageways in the first airfoil and one or more second feed passageways in the second airfoil. At least a first side selected from the pressure side of the first airfoil and the suction side of the second airfoil includes a first region with a local thinning or gap in the coating. Along the first side, the cooling passageway system includes means for locally cooling said first region.

Abstract: A turbine airfoil comprises a wall portion, a cooling channel, an impingement rib, impingement rib nozzles, turbulators and leading edge cooling holes. The wall portion comprises a leading edge, a trailing edge, an outer diameter end, and an inner diameter end. The cooling channel receives cooling air and extends through an interior of the wall portion between the inner diameter end and the outer diameter end. The impingement rib is positioned within the wall portion forward of the cooling channel and between the outer diameter end and the inner diameter end to define a peanut cavity. The impingement rib nozzles extend through the impingement rib for receiving cooling air from the cooling channel. The turbulators are positioned within the peanut cavity to locally influence the flow of the cooling air. The leading edge cooling holes discharge the cooling air from the peanut cavity to an exterior of the wall portion.

Abstract: A vane for use in a gas turbine engine has a platform connected to an airfoil. There is a cooling passage for supplying cooling air to the platform. A cooling chamber supplies cooling air to a plurality of cooling slots at the platform. The cooling slots have a non-uniform cross section.

Abstract: An airfoil assembly includes an airfoil with at least one cavity that is in communication with a source of cooling air. A baffle is disposed within that cavity and includes a plurality of openings for directing cooling air against the hot wall. A plurality of dividers extends between the baffle walls and the internal cavity to direct cooling air toward one of a leading edge chamber and a trailing edge chamber.

Abstract: A cooled airfoil includes an impingement rib having a multiple of openings which supply a cooling airflow from a cooling circuit flow path toward an airfoil leading edge. The multiple of openings are offset in the impingement rib opposite an outer airfoil wall which includes gill holes. Offsetting the multiple of openings opposite an outer airfoil wall which includes the gill holes focuses the cooling airflow across turbulators to increase the cooling airflow dwell time to increase the thermal transfer therefrom in higher temperature airfoil areas.

Abstract: A turbine airfoil comprises a wall portion, a cooling channel, an impingement rib, impingement rib nozzles, turbulators and leading edge cooling holes. The wall portion comprises a leading edge, a trailing edge, an outer diameter end, and an inner diameter end. The cooling channel receives cooling air and extends through an interior of the wall portion between the inner diameter end and the outer diameter end. The impingement rib is positioned within the wall portion forward of the cooling channel and between the outer diameter end and the inner diameter end to define a peanut cavity. The impingement rib nozzles extend through the impingement rib for receiving cooling air from the cooling channel. The turbulators are positioned within the peanut cavity to locally influence the flow of the cooling air. The leading edge cooling holes discharge the cooling air from the peanut cavity to an exterior of the wall portion.

Abstract: A component for use in a gas turbine engine is provided. The component has an airfoil portion with a plurality of internal cooling passages and a non-linear trailing edge. The component further has a non-linear array of teardrop shaped assemblies which form a plurality of injection slots for injecting a coolant fluid into a fluid passing over the airfoil portion. The teardrop shaped assemblies are designed to maximize thermal performance of the component by reducing a relative diffusion angle between the injected coolant flow and the streamline direction of the fluid passing over the airfoil portion.

Abstract: A vane cluster has a coated metallic substrate. The cluster includes a platform and a shroud. At least first and second airfoils extend between an outer face of the platform and an inner face of the shroud. Each airfoil has a pressure side and a suction side. The pressure side of the first airfoil faces the suction side of the second airfoil. The cluster includes a cooling passageway system including one or more first feed passageways in the first airfoil and one or more second feed passageways in the second airfoil. At least a first side selected from the pressure side of the first airfoil and the suction side of the second airfoil includes a first region with a local thinning or gap in the coating. Along the first side, the cooling passageway system includes means for locally cooling said first region.

Abstract: A component for use in a gas turbine engine is provided. The component has an airfoil portion with a plurality of internal cooling passages and a non-linear trailing edge. The component further has a non-linear array of teardrop shaped assemblies which form a plurality of injection slots for injecting a coolant fluid into a fluid passing over the airfoil portion. The teardrop shaped assemblies are designed to maximize thermal performance of the component by reducing a relative diffusion angle between the injected coolant flow and the streamline direction of the fluid passing over the airfoil portion.

Abstract: A turbine engine component, such as a blade or vane, has a system for cooling a trailing edge portion thereof. The system includes a plurality of rows of pedestals which vary in density along a span of the component. In a preferred embodiment of the present invention, the number of rows of pedestals increases as one moves along the span of the component from an inner diameter region to an outer diameter region.

Abstract: A turbine airfoil section having an internal cavity and a plurality of indentations on the inner surface of the internal cavity is described. The indentations provide enhanced heat transfer for cooling the internal cavity of an airfoil thereby improving the life of the airfoil and optimizing the efficiency of the engine by minimizing the amount of compressor bleed air required. Advantageously, this cooling scheme also does not restrict the cooling flow within the internal cavity. The indentations may have varying patterns and alternative geometric configurations.

Abstract: A turbine airfoil section having an internal cavity and a plurality of indentations on the inner surface of the internal cavity is described. The indentations provide enhanced heat transfer for cooling the internal cavity of an airfoil thereby improving the life of the airfoil and optimizing the efficiency of the engine by minimizing the amount of compressor bleed air required. Advantageously, this cooling scheme also does not restrict the cooling flow within the internal cavity. The indentations may have varying patterns and alternative geometric configurations.