# US 2019/0169997 A1 · Double wall turbine gas turbine engine blade cooling configuration

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- Number: US 2019/0169997 A1
- Document ID: US20190169997A1
- Status: Published application
- Kind: A1
- Application number: 15/832,478
- Filed: 2017-12-05
- Published: 2019-06-06
- Earliest claimed: 2017-12-05
- Assignees: United Technologies Corporation
- Applicants: United Technologies Corporation
- Inventors: Christopher Whitfield; Carey Clum; Dominic J. Mongillo, Jr.
- Web page: https://patentgrep.com/patent/US20190169997A1

## CPC

- F01D 5/187: {Convection cooling}
- F02C 9/18: by bleeding, bypassing or acting on variable working fluid interconnections between turbines or compressors or their stages {(F02C3/113 takes precedence)}
- F02C 3/04: having a turbine driving a compressor (power transmission arrangements F02C7/36; control of working fluid flow F02C9/16)
- F05D 2220/32: in gas turbines
- F05D 2240/35: Combustors or associated equipment
- F05D 2250/185: serpentine-like
- F05D 2260/201: by impingement of a fluid
- F05D 2260/202: by film cooling
- F05D 2260/204: by the use of microcircuits
- F05D 2260/2212: by creating turbulence
- Y02T 50/60: Efficient propulsion technologies, e.g. for aircraft
- F01D: NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES (machines or engines for liquids F03; non-positive displacement pumps F04D)
- F02C: GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS (construction of turbines F01D; jet-propulsion plants F02K; construction of compressors or fans F04; generating combustion products of high pressure or high velocity F23R; using gas turbines in compression refrigeration plants F25B11/00)
- F05D: INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- Y02T: CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION

## 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. 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 low turbulence flow region in the skin passage. The resupply hole has an exit at the skin passage and the exit has a diffuser.

## Claims

### Claim 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, 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 low turbulence flow region in the skin passage, wherein the resupply hole has an exit at the skin passage, the exit has a diffuser.

### Claim 2 (depends on 1)

The airfoil of claim 1, wherein the angle is in a range of 5°-45°.

### Claim 3 (depends on 1)

The airfoil of claim 1, 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.

### Claim 4 (depends on 3)

The airfoil of claim 3, wherein the passage height (H) is in a range of 0.010-0.200 inches (0.25-5.08 mm).

### Claim 5 (depends on 1)

The airfoil of claim 1, wherein the diffuser provides an aperture provided on an inner surface of the cold wall facing the hot wall, the aperture includes a length and a width, and a ratio of length to width in a range of 0.5:1 to 3:1.

### Claim 6 (depends on 5)

The airfoil of claim 5, wherein the diffuser has a bottom wall joined to laterally spaced apart side walls at intersections providing rounded corners.

### Claim 7 (depends on 5)

The airfoil of claim 5, wherein the diffuser includes multiple lobes having peaks and valleys extending in a direction of the centerline, the peaks and valleys providing scallops in the inner surface.

### Claim 8 (depends on 5)

The airfoil of claim 5, wherein the diffuser has a bottom wall joined to laterally spaced apart side walls at intersections, the bottom wall narrowing to a neck.

### Claim 9 (depends on 1)

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.

### Claim 10 (depends on 1)

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

### Claim 11 (depends on 1)

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

### Claim 12

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, 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 low turbulence flow region in the skin passage, wherein the resupply hole has an exit at the skin passage, the exit has a diffuser.

### Claim 13 (depends on 12)

The gas turbine engine of claim 12, wherein the angle is in a range of 5°-45°.

### Claim 14 (depends on 12)

The gas turbine engine of claim 12, 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).

### Claim 15 (depends on 12)

The gas turbine engine of claim 12, wherein the diffuser provides an aperture in an inner surface provided on an inner surface of the cold wall facing the hot wall, the aperture includes a length and a width, and a ratio of length to width in a range of 0.5:1 to 3:1.

### Claim 16 (depends on 15)

The gas turbine engine of claim 15, wherein the diffuser has a bottom wall joined to laterally spaced apart side walls at intersections providing rounded corners.

### Claim 17 (depends on 15)

The gas turbine engine of claim 15, wherein the diffuser includes multiple lobes having peaks and valleys extending in a direction of the centerline, the peaks and valleys providing scallops in the inner surface.

### Claim 18 (depends on 15)

The gas turbine engine of claim 15, wherein the diffuser has a bottom wall joined to laterally spaced apart side walls at intersections, the bottom wall narrowing to a neck.

### Claim 19 (depends on 12)

The gas turbine engine of claim 12, 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.

### Claim 20 (depends on 12)

The gas turbine engine of claim 12, wherein a film cooling hole extends from the skin passage to the exterior airfoil surface.