Forward flowing serpentine vane

Patent metadata

Field	Value
Publication number	US-10480328-B2
Application number	US-201715414103-A
Country	US
Kind code	B2
Filing date	Jan 24, 2017
Priority date	Jan 25, 2016
Publication date	Nov 19, 2019
Grant date	Nov 19, 2019

Abstract

A vane adapted for use in a gas turbine engine includes a forward following serpentine cooling path having a plurality of divider walls to direct cooling airflow. The serpentine cooling path optionally includes one or more bifurcating walls in the cooling path to split the airflow. The serpentine cooling path is in communication with an aft inlet in the vane and air passing through the serpentine cooling path exits a forward outlet in the vane into a forward rotor/stator cavity.

First claim

What is claimed is:

1. A vane adapted for use in a gas turbine engine, the vane comprising
an outer platform,
an inner platform, and
an airfoil that extends from the outer platform to the inner platform in a radial direction, the airfoil having a leading edge and a trailing edge and being formed to include a serpentine cooling passage with an inlet arranged through the outer platform at the trailing edge of the airfoil and an outlet arranged through the inner platform adjacent to the leading edge of the airfoil so that cooling air supplied to the inlet moves toward the leading edge of the airfoil during use in the gas turbine engine.

2. The vane of claim 1 , wherein the airfoil includes an outer skin and a plurality of interior divider walls that define at least an aft chamber arranged adjacent to the trailing edge of the airfoil, a forward chamber arranged adjacent to the leading edge of the airfoil, and a middle chamber arranged between the aft chamber and the forward chamber.

3. The vane of claim 2 , wherein the aft chamber is fluidly coupled to the middle chamber adjacent to the inner platform and the forward chamber is fluidly coupled to the middle chamber adjacent to the outer platform to provide the serpentine cooling passage.

4. The vane of claim 3 , wherein the bifurcation wall extends axially between a first divider wall separating the aft chamber from the middle chamber and a second divider wall separating the middle chamber from the forward chamber and the bifurcation wall is spaced apart radially from both the inner platform and the outer platform.

5. The vane of claim 2 , wherein the middle chamber further comprises a bifurcation wall that separates cooling air passing through the middle chamber into suction side and pressure side passages.

6. The vane of claim 5 , wherein the pressure side passage has smooth walls and the suction side passage has turbulator ribs to increase heat transfer to air that flows through the suction side passage.

7. The vane of claim 5 , wherein the pressure side passage has turbulator ribs of a first geometry and the suction side passage has turbulator ribs of a second geometry, wherein the first geometry is different than the second geometry.

8. The vane of claim 2 , wherein the aft chamber is fluidly coupled to the middle chamber adjacent to the inner platform by a first passageway that extends from the aft chamber to the middle chamber within the serpentine cooling passage and the forward chamber is fluidly coupled to the middle chamber adjacent to the outer platform by a second passageway that extends from the middle chamber to the forward chamber within the serpentine cooling passage to provide the serpentine cooling passage.

9. The vane of claim 1 , wherein the outlet arranged through the inner platform comprises a nozzle with a preswirl feature.

10. The vane of claim 1 , wherein the airfoil includes an outer skin and a plurality of interior walls that define at least an upper chamber arranged adjacent to the outer platform edge of the vane, a lower chamber arranged adjacent to inner platform of the vane, and a central chamber arranged between the upper chamber and the lower chamber.

11. The vane of claim 10 , wherein the upper chamber is fluidly coupled to the central chamber adjacent to the leading edge and the lower chamber is fluidly coupled to the central chamber adjacent to the trailing edge to provide the serpentine cooling passage.

12. A system for cooling stator cavities below a vane comprising
the vane having an outer platform,
an inner platform, and
an airfoil that extends from the outer platform to the inner platform in a radial direction, the airfoil having a leading edge and a trailing edge and being formed to include a serpentine cooling passage with an aft inlet arranged through the outer platform at the trailing edge of the airfoil, a forward inlet arranged through the outer platform adjacent to the forward edge of the airfoil, and an outlet arranged through the inner platform adjacent to the leading edge of the airfoil so that cooling air supplied to each inlet moves toward the leading edge of the airfoil during use in a gas turbine engine.

13. The system of claim 12 , wherein the outlet further comprises a nozzle with a preswirl feature.

14. The system of claim 12 , further comprising an aft outlet in the inner platform of the vane configured to direct a portion of the cooling air from the aft inlet into an aft stator cavity.

15. The system of claim 12 , wherein the serpentine cooling passage includes a plurality of divider walls and a bifurcation wall perpendicular to the plurality of divider walls.

16. The system of claim 15 , wherein the plurality of divider walls extend from the leading edge and the trailing edge of the airfoil and define at least an upper chamber arranged adjacent to the outer platform edge of the vane, a lower chamber arranged adjacent to inner platform of the vane, and a central chamber arranged between the upper chamber and the lower chamber.

17. The system of claim 15 , wherein the plurality divider walls extend from the outer platform and inner platform and define at least an aft chamber arranged adjacent to the trailing edge of the airfoil, a forward chamber arranged adjacent to the leading edge of the airfoil, and a middle chamber arranged between the aft chamber and the forward chamber.

18. The system of claim 17 wherein the bifurcation wall is arranged in the middle chamber and separates cooling air passing through the middle chamber into suction side and pressure side passages.

19. The system of claim 15 , wherein the serpentine cooling passage is defined entirely, radially between the inner platform and the outer platform.

20. The system of claim 12 , wherein the airfoil includes an outer skin and at least one interior divider wall extending radially between the inner platform and the outer platform and defining an aft chamber directly between the trailing edge of the airfoil and the first divider wall and the inlet of the serpentine cooling passage opens into the aft chamber.

Assignees

• Rolls Royce Corp
• Rolls Royce Nam Tech Inc

Inventors

• Weaver John Alan
• Riehle Bradford John

Classifications

• F05D2260/22141

  using fins or ribs · CPC title

• F05D2260/221

  Improvement of heat transfer · CPC title

• F05D2240/127

  Vortex generators, turbulators, or the like, for mixing · CPC title

• F05D2240/121

  related to the leading edge of a stator vane · CPC title

• F01D5/18 Primary

  Hollow blades, {i.e. blades with cooling or heating channels or cavities (structure of hollow blades in general F01D5/147)}; Heating, heat-insulating or cooling means on blades · CPC title