Double wall turbine gas turbine engine blade cooling configuration

Patent metadata

Field	Value
Publication number	US-10508555-B2
Application number	US-201715832407-A
Country	US
Kind code	B2
Filing date	Dec 5, 2017
Priority date	Dec 5, 2017
Publication date	Dec 17, 2019
Grant date	Dec 17, 2019

Abstract

An airfoil includes pressure and suction side walls that extend in a chord-wise direction between leading and trailing edges. The pressure and suction side walls extend in a radial direction to provide an exterior airfoil surface. A core cooling passage is arranged between the pressure and suction walls in a thickness direction and extends radially toward a tip. A skin passage is arranged in one of the pressure and suction side walls to form a hot side wall and a cold side wall. The hot side wall defines a portion of the exterior airfoil surface and the cold side wall defines a portion of the core passage. The core passage and the skin passage are configured to have a same direction of fluid flow. A resupply hole cluster includes first and second resupply holes that each fluidly interconnect the core and skin passages. The first and second resupply holes intersect one another and terminate in an exit at the skin passage.

First claim

What is claimed is:

1. An airfoil comprising:
pressure and suction side walls extending in a chord-wise direction between leading and
trailing edges, the pressure and suction side walls extending in a radial direction to provide an exterior airfoil surface, a core cooling passage is arranged between the pressure and suction walls in a thickness direction and extends radially toward a tip, a skin passage is arranged in one of the pressure and suction side walls to form a hot side wall and a cold side wall, the hot side wall defines a portion of the exterior airfoil surface, and the cold side wall defines a portion of the core passage, the core passage and the skin passage are configured to have a same direction of fluid flow, and a resupply hole cluster in the cold side wall including first and second resupply holes that each fluidly interconnect the core and skin passages, the first and second resupply holes intersect one another and terminate in an exit at the skin passage.

2. The airfoil of claim 1 , a centerline of each of the first and second resupply holes is arranged as an acute angle relative to the direction of fluid flow in the core passage and is configured to provide a low turbulence flow region in the skin passage.

3. The airfoil of claim 2 , wherein the angle is in a range of 60°-120°.

4. The airfoil of claim 1 , wherein the skin passage has an aspect ratio between 3:1≥H/W≥1:5, wherein H corresponds to a passage height and W corresponds to a passage width.

5. The airfoil of claim 4 , wherein the skin passage height (H) is in a range of 0.010-0.200 inches (0.25-5.08 mm).

6. The airfoil of claim 1 , comprising serpentine cooling passage having first, second and third cooling passages, the first and third cooling passages having a direction of fluid flow toward the tip, and the second cooling passage having a direction of fluid flow away from the tip, the core passage provided by one of the first, second and third cooling passages.

7. The airfoil of claim 1 , wherein a film cooling hole extends from the skin passage to the exterior airfoil surface.

8. The airfoil of claim 1 , wherein the airfoil is a turbine blade.

9. The airfoil of claim 1 , wherein a meter is provided at the intersection, the meter is configured to regulate a fluid flow to the exit, wherein the exit has a total exit area, and the meter has a meter area, a ratio of total exit area to meter area is 2:1 or greater.

10. The airfoil of claim 9 , wherein the ratio is greater than 5:1.

11. The airfoil of claim 10 , wherein the ratio is in a range of 5:1 to 10:1.

12. The airfoil of claim 1 , wherein the exit is provided with first and second exits respectively provided by the first and second resupply holes, the first and second exits discrete from one another.

13. The airfoil of claim 12 , wherein the first and second exits are one of oval shape, fanshaped, laidback fanshaped, and lobed diffusers.

14. The airfoil of claim 1 , wherein the exit is provided with first and second exits respectively provided by the first and second resupply holes, the first and second exits overlapping one another.

15. The airfoil of claim 14 , wherein the first and second exits each are provided by a diffuser.

16. The airfoil of claim 1 , wherein the first and second resupply holes each include a centerline arranged at an angle in a range of 60°-120°.

17. A gas turbine engine comprising:
a combustor section arranged fluidly between compressor and turbine sections; and
an airfoil arranged in the turbine section, the airfoil including pressure and suction side walls extending in a chord-wise direction between leading and trailing edges, the pressure and suction side walls extending in a radial direction to provide an exterior airfoil surface, a core cooling passage is arranged between the pressure and suction walls in a thickness direction and extends radially toward a tip, a skin passage is arranged in one of the pressure and suction side walls to form a hot side wall and a cold side wall, the hot side wall defines a portion of the exterior airfoil surface, and the cold side wall defines a portion of the core passage, the core passage and the skin passage are configured to receive a cooling fluid from the compressor section and have a same direction of fluid flow, and a resupply hole cluster in the cold side wall including first and second resupply holes that each fluidly interconnect the core and skin passages, the first and second resupply holes intersect one another and terminate in an exit at the skin passage.

18. The gas turbine engine of claim 17 , a centerline of the resupply hole is arranged at an acute angle relative to the direction of fluid flow in the core passage and is configured to provide a laminar flow region in the skin passage, wherein the angle is in a range of 60°-120°.

19. The gas turbine engine of claim 18 , wherein the skin passage has an aspect ratio between 3:1≥H/W≥1:5, wherein H corresponds to a passage height and W corresponds to a passage width, wherein the passage height (H) is in a range of 0.010-0.200 inches (0.25-5.08 mm).

20. The gas turbine engine of claim 17 , wherein a meter is provided at the intersection, the meter is configured to regulate a fluid flow to the exit, wherein the exit has a total exit area, and the meter has a meter area, a ratio of total exit area to meter area is greater than 5:1.

Assignees

• United Technologies Corp

Inventors

• Whitfield Christopher
• Clum Carey
• Mongillo Jr Dominic J

Classifications

• F05D2260/201

  by impingement of a fluid · CPC title

• F05D2260/202

  by film cooling · CPC title

• F05D2260/204

  by the use of microcircuits · CPC title

• F01D5/187 Primary

  Convection cooling · CPC title

• F02C3/04

  having a turbine driving a compressor (power transmission arrangements F02C7/36; control of working fluid flow F02C9/16) · CPC title