Double wall turbine gas turbine engine blade cooling configuration

Patent metadata

Field	Value
Publication number	US-10648345-B2
Application number	US-201715832467-A
Country	US
Kind code	B2
Filing date	Dec 5, 2017
Priority date	Dec 5, 2017
Publication date	May 12, 2020
Grant date	May 12, 2020

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 fluidly interconnects the core and skin passages. The resupply hole has a single inlet that is fluidly connected to multiple discrete outlet apertures. A centerline of the single inlet is arranged at an 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.

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 fluidly interconnects the core and skin passages, the resupply hole has a single inlet that is fluidly connected to multiple discrete outlet apertures, a centerline of the single inlet is arranged at an 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, wherein the angle is in a range of 5° -45°, wherein the skin passage has an aspect ratio that may vary 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).

2. The airfoil of claim 1 , wherein the inlet has an inlet hydraulic diameter, and the multiple discrete outlet apertures collectively provide an exit hydraulic diameter, the ratio of exit hydraulic diameter to inlet hydraulic diameter is 2:1 or greater.

3. The airfoil of claim 2 , wherein the ratio of exit hydraulic diameter to inlet hydraulic diameter is in a range of 5:1 to 10:1.

4. The airfoil of claim 2 , wherein the inlet hydraulic diameter is provided by a meter that defines the narrowest flow area of the resupply hole.

5. The airfoil of claim 1 , wherein the resupply hole includes first and second exit legs that terminated in first and second discrete outlet apertures.

6. The airfoil of claim 5 , wherein the first and second exit legs respectively have first and second centerlines that are each at an acute angle relative to the centerline of the inlet.

7. The airfoil of claim 5 , wherein the first and second discrete outlet apertures include a diffuser.

8. The airfoil of claim 1 , comprising a 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.

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

10. 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 fluidly interconnects the core and skin passages, the resupply hole has a single inlet that is fluidly connected to multiple discrete outlet apertures, a centerline of the single inlet is arranged at an 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, wherein the angle is in a range of 5° -45°, wherein the skin passage has an aspect ratio that may vary 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).

11. The gas turbine engine of claim 10 , wherein the inlet has an inlet hydraulic diameter, and the multiple discrete outlet apertures collectively provide an exit hydraulic diameter, the ratio of exit hydraulic diameter to inlet hydraulic diameter is 2:1 or greater.

12. The gas turbine engine of claim 11 , wherein the ratio of exit hydraulic diameter to inlet hydraulic diameter is in a range of 5:1 to 10:1, the inlet hydraulic diameter is provided by a meter that defines the narrowest flow area of the resupply hole.

13. The gas turbine engine of claim 12 , wherein the resupply hole includes first and second exit legs that terminated in first and second discrete outlet apertures.

14. The gas turbine engine of claim 13 , wherein the first and second exit legs respectively have first and second centerlines that are each at an acute angle relative to the centerline of the inlet.

15. The gas turbine engine of claim 10 , wherein the airfoil is a turbine blade, comprising a 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.

Assignees

• United Technologies Corp

Inventors

• Whitfield Christopher
• Clum Carey
• Mongillo Jr Dominic J

Classifications

• F05D2260/2212

  by creating turbulence · CPC title

• F05D2260/202

  by film cooling · CPC title

• F05D2220/32

  in gas turbines · CPC title

• F02C3/04

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

• F02C9/18

  by bleeding, bypassing or acting on variable working fluid interconnections between turbines or compressors or their stages {(F02C3/113 takes precedence)} · CPC title