Turbine blade

Patent metadata

Field	Value
Publication number	US-2017107828-A1
Application number	US-201514884100-A
Country	US
Kind code	A1
Filing date	Oct 15, 2015
Priority date	Oct 15, 2015
Publication date	Apr 20, 2017
Grant date	—

Abstract

An airfoil for a gas turbine engine comprises a cooling circuit defined within the airfoil providing a flow of cooling fluid within the airfoil. The cooling circuit exhausts the flow of cooling fluid out a slot opening comprising an airfoil element. The slot opening further defines an acceleration zone and a deceleration zone to meter the flow of cooling fluid within the airfoil.

First claim

What is claimed is:

1 . An airfoil for a gas turbine engine, the airfoil comprising:
an outer surface defining a pressure side and a suction side extending axially between a leading edge and a trailing edge and extending radially between a root and a tip, and the trailing edge having a slot opening; a cooling circuit located within the airfoil and comprising a cooling passage fluidly coupled to the cooling air inlet passage and extending from the root toward the tip and terminating in an aft turn fluidly coupled to the slot opening; and an airfoil element located within the cooling passage downstream of the turn and upstream of the slot opening and forming an acceleration zone in the cooling passage along an upstream portion of the airfoil element and a deceleration zone in the cooling passage along a downstream portion of the airfoil element.

2 . The airfoil according to claim 1 wherein the airfoil element has an increasing cross-sectional area, which in combination with a first surrounding portion of the cooling passage forms a reducing cross-sectional area of the cooling passage to define the acceleration zone, and a decreasing cross-sectional area, which in combination with a second surrounding portion of the cooling passage forms an increasing cross-sectional area of the cooling passage to define the deceleration zone.

3 . The airfoil according to claim 2 wherein a third portion of the cooling passage immediately upstream of the airfoil element has a decreasing cross-sectional area to further define the acceleration zone.

4 . The airfoil according to claim 3 wherein the third portion of the cooling passage overlaps in a flow direction with the increasing cross-sectional area of the airfoil element.

5 . The airfoil according to claim 2 wherein a fourth portion of the cooling passage immediately downstream of the airfoil element has an increasing cross-sectional area to further define the deceleration zone.

6 . The airfoil according to claim 5 wherein the fourth portion of the cooling passage overlaps in a flow direction with the decreasing cross-sectional area of the airfoil element.

7 . The airfoil according to claim 1 wherein the slot opening further extends along a portion of the tip.

8 . The airfoil according to claim 1 further comprising turbulators located within the cooling passage.

9 . The airfoil according to claim 8 wherein the turbulators are located through the turn.

10 . The airfoil according to claim 1 wherein the airfoil element extends between the pressure and suction sides.

11 . A blade for a gas turbine engine having a turbine rotor disk, the blade comprising:
a dovetail having at least one cooling air inlet passage and configured to mount to the turbine rotor disk; an airfoil extending radially from the dovetail and having an outer surface defining a pressure side and a suction side extending axially between a leading edge and a trailing edge and extending radially between a root and a tip, with the root being adjacent the dovetail, and the trailing edge having a slot opening; a cooling circuit located within the airfoil and comprising a cooling passage fluidly coupled to the cooling air inlet passage and having multiple passes extending relatively between the root and the tip, with the multiple passes in a fore-to-aft serpentine arrangement, with the aft-most one of the multiple passes terminating in an aft turn fluidly coupled to the slot opening, and the cooling passage having an exit nozzle formed by a converging portion defining an acceleration zone and a diverging portion defining a deceleration zone, which are separated by a minimum cross-section area to define a choke, with the diverging portion located adjacent the slot opening; and an airfoil element located within the nozzle and extending between the pressure side and suction side.

12 . The blade according to claim 11 wherein there are three passes, with the fore-most and aft-most passes extending in a root-to-tip direction and the other pass extending in a tip-to-root direction.

13 . The blade according to claim 12 further comprising turbulators located within at least the aft-most pass.

14 . The blade according to claim 13 wherein the turbulators are located through the turn.

15 . The blade according to claim 14 wherein the turbulators are located in all of the three passes.

16 . The blade according to claim 14 wherein the slot opening further extends along a portion of the tip.

17 . The blade according to claim 17 wherein the airfoil element has a maximum thickness and the airfoil element is located within the nozzle such that the maximum thickness is aligned with the choke.

18 . The blade according to claim 11 wherein the slot opening further extends along a portion of the tip.

19 . The blade according to claim 11 wherein the airfoil element extends in a chord-wise direction aft of the deceleration zone and through the slot opening.

20 . A blade for a gas turbine engine comprising an airfoil having a pressure side and a suction side extending chord-wise between a leading edge and a trailing edge and extending span-wise between a root and a tip, a trailing edge cooling circuit located within the airfoil near the trailing edge and terminating in an aft turn fluidly coupled to a trailing edge slot opening, with an airfoil element located within the cooling circuit downstream of the turn and upstream of the slot opening and forming an acceleration zone in the cooling circuit along an upstream portion of the airfoil element and a deceleration zone in the cooling circuit along a downstream portion of the airfoil element.

Assignees

• Gen Electric

Inventors

• Krumanaker Matthew Lee
• Dooley Weston Nolan
• Brassfield Steven Robert

Classifications

• F01D5/147

  Construction, i.e. structural features, e.g. of weight-saving hollow blades (F01D5/148, F01D5/16 and F01D5/20 take precedence; blade shape F01D5/141; blades with cooling or heating channels or cavities F01D5/18; heating, heat-insulating or cooling means on blades F01D5/18) · CPC title

• F05D2260/22141

  using fins or ribs · CPC title

• F05D2240/307

  related to the tip of a rotor blade · 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

• F05D2260/20

  Heat transfer, e.g. cooling · CPC title