Radial-axial cooling slots
An engine component for a gas turbine engine may include a radial channel and an axial channel disposed in the engine component. The radial channel may be configured to direct a cooling air in a radial direction. The axial channel may be configured to direct the cooling air in a direction substantially perpendicular to the radial direction. A cross-section area of the axial channel is greater than a cross-section area of the radial channel, such that the radial channel remains the metering channel, independent of the relative location due to radial movement of the engine component with respect to an adjacent engine component.
1. An engine component for a gas turbine engine comprising:
a radial channel disposed in the engine component, the radial channel configured to direct a cooling air in a radial direction; and
an axial channel disposed in the engine component, the axial channel configured to direct the cooling air in a direction substantially perpendicular to the radial direction,
wherein a cross-section area of the axial channel is greater than a cross-section area of the radial channel,
wherein the engine component comprises an annular feature defined by at least a proximal surface, a distal surface, an aft side and a forward side;
the radial channel is disposed in at least one of the forward side or the aft side and extends between the proximal surface and the distal surface; and
the axial channel is disposed in the distal surface and extends from at least one of the forward side or the aft side and circumferentially in line with the radial channel.
2. The engine component of claim 1 , wherein the radial channel is disposed on the aft side and the axial channel extends from the aft side.
3. The engine component of claim 1 , wherein the radial channel is disposed on the forward side and the axial channel extends from the forward side.
4. The engine component of claim 1 , wherein the radial channel and the axial channel are configured to direct the cooling air between a proximal side of the engine component and a distal side of the engine component for cooling the engine component.
5. The engine component of claim 4 , wherein the engine component is configured to receive the cooling air from an aperture disposed in a rotor disk leg, the rotor disk leg being located radially inward from the engine component.
6. The engine component of claim 1 , wherein the engine component is configured to be coupled between a forward rotor disk and an aft rotor disk.
7. The engine component of claim 1 , wherein the engine component comprises knife edges extending from the distal surface, the knife edges configured to interface with a proximal surface of a vane platform.
8. The engine component of claim 1 , wherein the engine component comprises a nickel-based alloy.
9. A gas turbine engine comprising:
a compressor section;
a combustor section;
a turbine section;
an aft blade disk;
a forward blade disk; and
an engine component coupled between the aft blade disk and the forward blade disk comprising:
a radial channel disposed in the engine component, the radial channel configured to direct a cooling air in a radial direction; and
an axial channel disposed in the engine component, the axial channel configured to direct the cooling air in a direction substantially perpendicular to the radial direction,
wherein a cross-section area of the axial channel is greater than a cross-section area of the radial channel,
wherein the engine component comprises an annular feature defined by at least a proximal surface, a distal surface, an aft side and a forward side;
the radial channel is disposed in at least one of the forward side or the aft side and extends between the proximal surface and the distal surface; and
the axial channel is disposed in the distal surface and extends from at least one of the forward side or the aft side and circumferentially in line with the radial channel.
10. The gas turbine engine of claim 9 , wherein the radial channel is disposed on the aft side and the axial channel extends from the aft side.
11. The gas turbine engine of claim 9 , wherein the radial channel is disposed on the forward side and the axial channel extends from the forward side.
12. The gas turbine engine of claim 9 , wherein the radial channel and the axial channel are configured to direct the cooling air between a proximal side of the engine component and a distal side of the engine component for cooling the engine component.
13. The gas turbine engine of claim 12 , wherein the engine component is configured to receive the cooling air from an aperture disposed in a rotor disk leg, the rotor disk leg being located radially inward from the engine component.
14. The gas turbine engine of claim 9 , wherein the engine component is configured to be coupled between a forward rotor disk and an aft rotor disk.
15. The gas turbine engine of claim 9 , wherein the engine component comprises knife edges extending from the distal surface, the knife edges configured to interface with a proximal surface of a vane platform.
16. A method of manufacturing an engine component for a gas turbine engine comprising:
forming a radial channel in a side surface of the engine component, the radial channel extending between a proximal surface and a distal surface; and
forming an axial channel in an axially extending surface of the engine component, the axial channel extending from at least one of the forward side or the aft side;
wherein the forming the radial channel and the forming the axial channel provides the radial channel having a cross-section area which is less than a cross-section area of the axial channel.
17. The method of claim 16 , wherein the forming the radial channel is performed by milling the at least one of the forward side or the aft side of the engine component.
18. The method of claim 16 , wherein the forming the radial channel and the forming the axial channel provides the axial channel circumferentially in line with the radial channel.