# US 12,215,601 B2 · Air foil with staggered cooling hole configuration

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- Number: US 12,215,601 B2
- Document ID: US12215601B2
- Status: Granted
- Kind: B2
- Application number: 18/170,810
- Filed: 2023-02-17
- Published: 2024-08-22
- Granted: 2025-02-04
- Earliest claimed: 2023-02-17
- Assignees: Rtx Corporation
- Applicants: RTX Corporation
- Inventors: Jaime G. Ghigliotty Rosado; Brandon W. Spangler; Tracy A. Propheter-Hinckley
- Web page: https://patentgrep.com/patent/US12215601B2

## CPC

- F01D 5/186: {Film cooling (F01D5/187 takes precedence)}
- F01D 5/187: {Convection cooling}
- F05D 2240/121: related to the leading edge of a stator vane
- F05D 2240/301: Cross-sectional characteristics
- F05D 2240/303: related to the leading edge of a rotor blade
- F05D 2260/201: by impingement of a fluid
- F05D 2260/202: by film cooling
- F01D: NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES (machines or engines for liquids F03; non-positive displacement pumps F04D)
- F05D: INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS

## Abstract

A turbine blade for a gas turbine engine, including: an airfoil, the having a leading edge, a pressure side, a suction side and a trailing edge; a plurality of internal cooling cavities including a leading edge cavity, a leading edge feed passage, pressure side cooling passages, suction side cooling passages and main body cavities; the leading edge cavity extending towards the suction side; a first crossover row of cooling passages providing fluid communication between the leading edge cavity and the leading edge feed passage; and a second crossover row of cooling passages providing fluid communication between the leading edge cavity and the leading edge feed passage, a centerline of the first crossover row of cooling passages is located closer to the pressure side than a centerline of the second crossover row of cooling passages and the centerline of the second crossover row of cooling passages is located closer to the suction side than the centerline of the first crossover row of cooling passages, and wherein the second crossover row of cooling passages are radially staggered relative to the first crossover row of cooling passages.

## Claims

### Claim 1

A turbine blade for a gas turbine engine, comprising:
An airfoil, the airfoil having a leading edge, a pressure side, a suction side and a trailing edge;
a plurality of internal cooling cavities including a leading edge cavity, a leading edge feed passage, pressure side cooling passages, suction side cooling passages and main body cavities; the leading edge cavity extending towards the suction side;
a first crossover row of cooling passages providing fluid communication between the leading edge cavity and the leading edge feed passage; and
a second crossover row of cooling passages providing fluid communication between the leading edge cavity and the leading edge feed passage, a centerline of the first crossover row of cooling passages is located closer to the pressure side than a centerline of the second crossover row of cooling passages and the centerline of the second crossover row of cooling passages is located closer to the suction side than the centerline of the first crossover row of cooling passages, and wherein the second crossover row of cooling passages are radially staggered relative to the first crossover row of cooling passages and the centerline of the first crossover row of cooling passages and the centerline of the second crossover row of cooling passages intersects the leading edge cavity at a point forward of a line parallel to a pull angle or edge of the leading edge feed passage and the centerline of the first crossover row of cooling passages and the centerline of the second crossover row of cooling passages intersects a vertex of the leading edge feed passage.

### Claim 2 (depends on 1)

The turbine blade according to claim 1, wherein the first crossover row of cooling passages and the second crossover row of cooling passages are angled with respect to a horizontal line extending between the leading edge cavity and the leading edge feed passage.

### Claim 3 (depends on 1)

The turbine blade according to claim 1, wherein the leading edge cavity proximate to the suction side is provided with an impingement cooling benefit from the second crossover row of cooling passages.

### Claim 4 (depends on 1)

The turbine blade according to claim 1, wherein the centerline of the first crossover row of cooling passages and the centerline of the second crossover row of cooling passages are each aligned with an angle gamma (γ) with respect to a horizontal line extending from the vertex of the leading edge feed passage to the vertex of the leading edge cavity, wherein the angle gamma (γ) of the first crossover row of cooling passages is less than or equal to a pull angle alpha (α) of a rib for forming the first crossover row of cooling passages, the pull angle alpha (α) being relative to the horizontal line extending from the vertex of the leading edge feed passage to the vertex of the leading edge cavity and the angle gamma (γ) of the second crossover row of cooling passages is less than or equal to a pull angle beta (β) of a rib for forming the second crossover row of cooling passages, the pull angle beta (β) being relative to the horizontal line extending from the vertex of the leading edge feed passage to the vertex of the leading edge cavity.

### Claim 5 (depends on 1)

The turbine blade according to claim 1, wherein the first crossover row of cooling passages and the second crossover row of cooling passages taper into the leading edge feed passage.

### Claim 6 (depends on 5)

The turbine blade according to claim 5, wherein at least one of the first crossover row of cooling passages and the second crossover row of cooling passages do not extend all the way to an exterior wall of the airfoil.

### Claim 7

A gas turbine engine comprising:
a compressor section;
a combustor fluidly connected to the compressor section;
a turbine section fluidly connected to the combustor, the turbine section comprising:
a high pressure turbine coupled to a high pressure compressor of the compressor section via a shaft;
a low pressure turbine; and
wherein the high pressure turbine includes a turbine disk with a plurality of turbine blades secured thereto each of the plurality of turbine blades, comprising:
an airfoil, the airfoil having a leading edge, a pressure side, a suction side and a trailing edge;
a plurality of internal cooling cavities including a leading edge cavity, a leading edge feed passage, pressure side cooling passages, suction side cooling passages and main body cavities; the leading edge cavity extending towards the suction side;
a first crossover row of cooling passages providing fluid communication between the leading edge cavity and the leading edge feed passage; and
a second crossover row of cooling passages providing fluid communication between the leading edge cavity and the leading edge feed passage, a centerline of the first crossover row of cooling passages is located closer to the pressure side than a centerline of the second crossover row of cooling passages and the second crossover row of cooling passages is located closer to the suction side than the centerline of the first crossover row of cooling passages, and wherein the second crossover row of cooling passages are radially staggered relative to the first crossover row of cooling passages and the centerline of the first crossover row of cooling passages and the centerline of the second crossover row of cooling passages intersects the leading edge cavity at a point forward of a line parallel to a pull angle or edge of the leading edge feed passage and the centerline of the first crossover row of cooling passages and the centerline of the second crossover row of cooling passages intersects a vertex of the leading edge feed passage.

### Claim 8 (depends on 7)

The gas turbine engine as in claim 7, wherein the first crossover row of cooling passages and the second crossover row of cooling passages are angled with respect to a horizontal line extending between the leading edge cavity and the leading edge feed passage.

### Claim 9 (depends on 7)

The gas turbine engine as in claim 7, wherein the leading edge cavity proximate to the suction side is provided with an impingement cooling benefit from the second crossover row of cooling passages.

### Claim 10 (depends on 7)

The gas turbine engine as in claim 7, wherein the first crossover row of cooling passages and the second crossover row of cooling passages taper into the leading edge feed passage.

### Claim 11 (depends on 10)

The gas turbine engine as in claim 10, wherein at least one of the first crossover row of cooling passages and the second crossover row of cooling passages do not extend all the way to an exterior wall of the airfoil.

### Claim 12

A method for forming an airfoil of a turbine blade, comprising:
forming a plurality of internal cooling cavities in the airfoil, the plurality of internal cooling cavities including a leading edge cavity, a leading edge feed passage, pressure side cooling passages, suction side cooling passages and main body cavities; the leading edge cavity extending towards the suction side, the airfoil having a leading edge, a pressure side, a suction side and a trailing edge;
forming a first crossover row of cooling passages providing fluid communication between the leading edge cavity and the leading edge feed passage; and
forming a second crossover row of cooling passages providing fluid communication between the leading edge cavity and the leading edge feed passage, a centerline of the first crossover row of cooling passages is located closer to the pressure side than a centerline of the second crossover row of cooling passages and the centerline of the second crossover row of cooling passages is located closer to the suction side than the centerline of the first crossover row of cooling passages, and wherein the second crossover row of cooling passages are radially staggered relative to the first crossover row of cooling passages and the centerline of the first crossover row of cooling passages and the centerline of the second crossover row of cooling passages intersects the leading edge cavity at a point forward of a line parallel to a pull angle or edge of the leading edge feed passage and the centerline of the first crossover row of cooling passages and the centerline of the second crossover row of cooling passages intersects a vertex of the leading edge feed passage.

### Claim 13 (depends on 12)

The method of claim 12, wherein the first crossover row of cooling passages and the second crossover row of cooling passages are angled with respect to a horizontal line extending between the leading edge cavity and the leading edge feed passage.

### Claim 14 (depends on 12)

The method of claim 12, wherein the leading edge cavity proximate to the suction side is provided with an impingement cooling benefit from the second crossover row of cooling passages.

### Claim 15 (depends on 12)

The method of claim 12, wherein the first crossover row of cooling passages and the second crossover row of cooling passages taper into the leading edge feed passage.