Axial flowing cooling passages for gas turbine engine components
A gas turbine engine component is provided. The gas turbine engine component comprises a main body and a leading edge cooling passage defined within the main body. The main body has a leading edge and a leading edge wall including an elongated transition portion extending between the leading edge and a proximate flowpath surface of the main body. The leading edge cooling passage comprises an axial flow cooling passage defined within the main body and adjacent to the leading edge wall and has a leading edge periphery that generally conforms to the elongated transition portion of the leading edge wall.
1. A gas turbine engine component comprising:
a main body having a leading edge and a leading edge wall including an elongated transition portion extending between the leading edge and a proximate flowpath surface of the main body and a cooling hole; and
a leading edge cooling passage comprising an axial flow cooling passage defined within the main body, the leading edge cooling passage adjacent to the leading edge wall and having a leading edge periphery that generally conforms to the elongated transition portion of the leading edge wall, wherein the leading edge periphery comprises an ellipse, wherein the elongated transition portion comprises an ellipse.
2. The gas turbine engine component of claim 1 , wherein the leading edge cooling passage has a leading edge cooling passage inlet aft of a leading edge cooling passage outlet, the leading edge cooling passage outlet communicating with the cooling hole, wherein a cooling flow through the gas turbine engine component comprises an axial cooling flow in a forward direction toward the leading edge wall.
3. The gas turbine engine component of claim 1 , wherein the cooling hole comprises a film cooling hole that is disposed in the elongated transition portion of the leading edge wall.
4. The gas turbine engine component of claim 3 , wherein the leading edge cooling passage outlet communicating with the film cooling hole provides film cooling of the leading edge portion of the main body.
5. The gas turbine engine component of claim 4 , wherein convective cooling occurs within the leading edge cooling passage, in addition to the film cooling.
6. The gas turbine engine component of claim 1 , wherein the leading edge cooling passage has a leading edge cooling passage inlet forward of a leading edge cooling passage outlet, wherein a cooling flow through the gas turbine engine component comprises an axial cooling flow in an aft direction away from the leading edge wall.
7. The gas turbine engine component of claim 1 , wherein the leading edge periphery is configured to substantially maintain a constant distance between the leading edge wall and the leading edge cooling passage.
8. The gas turbine engine component of claim 1 , wherein the main body and leading edge cooling passage are that of a blade outer air seal (BOAS) segment comprising the gas turbine engine component.
9. A blade outer air seal segment for a gas turbine engine comprising:
a main body that extends generally axially, with respect to a central axis, from a leading edge portion of the main body to a trailing edge portion of the main body, wherein the leading edge portion includes a leading edge wall having an elongated transition portion and a cooling hole; and
a leading edge cooling passage comprising an axial flow cooling passage defined within the main body, the leading edge cooling passage adjacent to the leading edge wall and having a leading edge periphery that generally conforms to the elongated transition portion of the leading edge wall, wherein the leading edge periphery comprises an ellipse, wherein the elongated transition portion comprises an ellipse.
10. The blade outer air seal segment of claim 9 , wherein the leading edge cooling passage has a leading edge cooling passage inlet aft of a leading edge cooling passage outlet, the leading edge cooling passage outlet communicating with the cooling hole and wherein a cooling flow through the blade outer air seal segment comprises an axial cooling flow in a forward direction toward the leading edge wall.
11. The blade outer air seal segment of claim 10 , wherein the cooling hole comprises a film cooling hole that is disposed in the elongated transition portion of the leading edge wall.
12. The blade outer air seal segment of claim 10 , wherein the leading edge cooling passage outlet communicating with the cooling hole provides film cooling of the leading edge portion of the main body.
13. The blade outer air seal segment of claim 12 , wherein convective cooling occurs within the blade outer air seal segment, in addition to the film cooling.
14. The blade outer air seal segment of claim 9 , wherein the leading edge cooling passage has a leading edge cooling passage inlet forward of a leading edge cooling passage outlet, wherein a cooling flow through a gas turbine engine component comprises an axial cooling flow in an aft direction away from the leading edge wall.
15. A gas turbine engine comprising:
a blade stage; and
a circumferential array of blade outer air seal segments in the blade stage, a blade outer air seal (BOAS) segment comprising:
a main body that extends generally axially with respect to a central axis, from a leading edge portion of the main body to a trailing edge portion of the main body, wherein the leading edge portion includes a film cooling hole, a leading edge and a leading edge wall having an elongated transition portion extending between the leading edge and an inner diameter flowpath surface of the main body; and
a leading edge cooling passage comprising an axial flow cooling passage defined within the main body, the leading edge cooling passage adjacent to the leading edge wall and having a leading edge periphery that generally conforms to the elongated transition portion of the leading edge wall, wherein the leading edge periphery comprises an ellipse, wherein the elongated transition portion comprises an ellipse.
16. The gas turbine engine of claim 15 , wherein the leading edge cooling passage has a leading edge cooling passage inlet aft of a leading edge cooling passage outlet, the leading edge cooling passage outlet communicating with the film cooling hole and wherein a cooling flow through the BOAS segment comprises an axial cooling flow in a forward direction toward the leading edge wall.
17. The gas turbine engine of claim 15 , wherein the film cooling hole is disposed in the elongated transition portion of the leading edge wall.
18. The gas turbine engine of claim 16 , wherein the leading edge cooling passage outlet communicating with the film cooling hole provides film cooling of the leading edge portion of the main body.
19. The gas turbine engine of claim 18 , wherein convective cooling occurs within the BOAS segment, in addition to the film cooling.
20. The gas turbine engine of claim 15 , wherein the leading edge cooling passage has a leading edge cooling passage inlet forward of a leading edge cooling passage outlet, wherein a cooling flow through a gas turbine engine component comprises an axial cooling flow in an aft direction away from the leading edge wall.