Film cooling passage with multidimensional diffusion
A film-cooled component for a gas turbine engine includes a first surface of the component located at a gas path of a gas turbine engine, a second surface of the component defining a component passage, and a cooling airflow passage extending from the second surface to the first surface to convey a cooling airflow from the passage and emit the cooling airflow at the first surface. The cooling airflow passage is curvilinearly diffused in at least two directions relative to a local gas flow direction in the gas path. A method of cooling a component includes flowing a cooling airflow into an internal component passage of the turbine component, conveying the cooling airflow through a cooling airflow passage, diffusing the cooling airflow in at least two directions along the airflow passage, and emitting the cooling airflow at a gas path surface to cool the gas path surface of the component.
1. A film-cooled component for a gas turbine engine, comprising:
a first surface of the component located at a gas path of the gas turbine engine;
a second surface of the component defining a component passage; and
a cooling airflow passage extending from the second surface to the first surface to convey a cooling airflow from the passage and emit the cooling airflow at the first surface, the cooling airflow passage diffusing in both a streamwise direction relative to a local hot gas flow in the gas path and a lateral direction perpendicular to the streamwise direction;
wherein the cooling airflow passage is defined along the lateral direction by a first passage surface defined by a first radius and a second passage surface defined by a second radius offset in the lateral direction from the first radius;
wherein the second radius is greater than the first radius;
wherein the cooling passage includes:
a non-circular passage inlet at the second surface of the component with an inlet cross-sectional major axis in the second surface perpendicular to the streamwise direction; and
a passage outlet at the first surface of the component with an outlet cross-sectional major axis in the first surface angularly offset from the inlet cross-sectional major axis and from the streamwise direction, the passage outlet one of oval or elliptical.
2. The film-cooled component of claim 1 , wherein the first surface is an external surface of the component and the second surface is an internal surface of the component.
3. The film-cooled component of claim 1 wherein:
the first radius is defined in a first plane at a first angle ranging from 15 degrees to 90 degrees to a gas flow direction; and
the second radius is defined in a second plane angularly offset from both the first plane and the gas flow direction.
4. The film-cooled component of claim 3 , wherein the second plane is angularly offset from the first plane in the range of 0 to 50 degrees.
5. The film-cooled component of claim 1 , wherein elongation of the cooling airflow passage increases continuously with distance from the second surface.
6. The film-cooled component of claim 1 , wherein the component is a turbine vane.
7. The film cooled component of claim 1 , wherein the component is formed via additive manufacturing.
8. A gas turbine engine, comprising:
a combustor; and
a turbine disposed in fluid communication with the combustor, the turbine having a plurality of turbine components, at least one turbine component of the plurality of turbine components including:
a first surface of the at least one turbine component located at a gas path of the gas turbine engine;
a second surface of the turbine component defining a component passage; and
a cooling airflow passage extending from the second surface to the first surface to convey a cooling airflow from the component passage and emit the cooling airflow at the first surface, the cooling airflow passage diffusing in both a streamwise direction relative to a local hot gas flow in the gas path and a lateral direction perpendicular to the streamwise direction as the cooling airflow passage extends from the second surface to the first surface;
wherein the cooling airflow passage is defined along the lateral direction by a first passage surface defined by a first radius and a second passage surface defined by a second radius offset in the lateral direction from the first radius;
wherein the second radius is greater than the first radius;
wherein the cooling passage includes:
a non-circular passage inlet at the second surface of the component with an inlet cross-sectional major axis in the second surface perpendicular to the streamwise direction; and
a passage outlet at the first surface of the component with an outlet cross-sectional major axis in the first surface angularly offset from the inlet cross-sectional major axis and from the streamwise direction, the passage outlet one of oval or elliptical.
9. The gas turbine engine of claim 8 , wherein the first surface is an external surface of the component and the second surface is an internal surface of the component.
10. The gas turbine engine of claim 8 , wherein:
the first radius is defined in a first plane at a first angle ranging from 15 degrees to 90 degrees to a gas flow direction; and
the second radius is defined in a second plane angularly offset from both the first plane and the gas flow direction.
11. The gas turbine engine of claim 10 , wherein the angular offset is in the range of 0 to 50 degrees.
12. The gas turbine engine of claim 8 , wherein a degree of diffusion of the cooling airflow passage increases continuously with distance from the second surface.
13. The gas turbine engine of claim 8 , wherein the turbine component is formed via additive manufacturing.