# US 10,370,976 B2 · Directional cooling arrangement for airfoils

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- Number: US 10,370,976 B2
- Document ID: US10370976B2
- Status: Granted
- Kind: B2
- Application number: 15/679,442
- Filed: 2017-08-17
- Published: 2019-02-21
- Granted: 2019-08-06
- Earliest claimed: 2017-08-17
- Assignees: United Technologies Corporation
- Applicants: United Technologies Corporation
- Inventors: San Quach; Alberto A. Mateo; Jeffrey T. Morton; Jason Shenny; Gregory Anselmi
- Web page: https://patentgrep.com/patent/US10370976B2

## CPC

- F01D 5/186: {Film cooling (F01D5/187 takes precedence)}
- F05D 2220/32: in gas turbines
- F05D 2240/303: related to the leading edge of a rotor blade
- F05D 2240/304: related to the trailing edge of a rotor blade
- F05D 2250/52: Outlet
- F05D 2260/202: by film cooling
- 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)
- 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 according to an example of the present disclosure includes, among other things, an internal wall and an external wall. The external wall defines pressure and suction sides between a leading edge and a trailing edge, and the airfoil section defines a mean camber line that extends between the leading and trailing edges to bisect a thickness of the airfoil section. A first cavity and a second cavity are separated by the internal wall. The second cavity is bounded by the external wall at the leading edge. At least one crossover passage within the internal wall connects the first cavity to the second cavity. The crossover passage defines a passage axis. The passage axis defines a passage angle with respect to the mean camber line such that the passage axis extends transversely from the mean camber line to intersect a surface of the second cavity.

## Claims

### Claim 1

An airfoil, comprising:
an airfoil section having an internal wall and an external wall, the external wall defining pressure and suction sides between a leading edge and a trailing edge, and the airfoil section defining a mean camber line extending between the leading and trailing edges to bisect a thickness of the airfoil section;
a first cavity and a second cavity separated by the internal wall, the second cavity bounded by the external wall at the leading edge;
at least one crossover passage within the internal wall that connects the first cavity to the second cavity; and
wherein the at least one crossover passage defines a passage axis, the passage axis defining a passage angle with respect to the mean camber line such that the passage axis extends transversely from the mean camber line to intersect a surface of the second cavity.

### Claim 2 (depends on 1)

The airfoil as recited in claim 1, wherein the passage angle is between 45 degrees and 75 degrees.

### Claim 3 (depends on 2)

The airfoil as recited in claim 2, wherein the passage axis intersects the suction side.

### Claim 4 (depends on 2)

The airfoil as recited in claim 2, wherein the passage axis intersects the pressure side.

### Claim 5 (depends on 1)

The airfoil as recited in claim 1, wherein the airfoil section extends in a spanwise direction from a 0% span position to a 100% span position, and an outlet of the at least one crossover passage is defined at a span position that is between an 80% span position and the 100% span position.

### Claim 6 (depends on 5)

The airfoil as recited in claim 5, wherein the passage angle is between 15 degrees and 90 degrees.

### Claim 7 (depends on 1)

The airfoil as recited in claim 1, comprising a plurality of film cooling passages within the external wall and extending from a surface of the second cavity, wherein the passage axis intersects the external wall at an intersection point, and the second cavity is free of any film cooling passages aft of the intersection point.

### Claim 8 (depends on 7)

The airfoil as recited in claim 7, wherein the mean camber line intersects the at least one crossover passage.

### Claim 9 (depends on 8)

The airfoil as recited in claim 8, wherein the passage axis intersects the suction side.

### Claim 10 (depends on 7)

The airfoil as recited in claim 7, wherein the at least one crossover passage extends between an inlet and an outlet, the outlet spaced apart from the external wall.

### Claim 11 (depends on 1)

The airfoil as recited in claim 1, wherein the airfoil is a turbine blade.

### Claim 12

A casting core for an airfoil, comprising:
a first portion corresponding to a leading edge cavity of an airfoil and a second portion corresponding to a second cavity of the airfoil, the airfoil defining a mean camber line that bisects a thickness of the airfoil between pressure and suction sides of the airfoil;
a plurality of connectors coupling the first portion and the second portion, the plurality of connectors corresponding to a plurality of crossover passages of the airfoil, each of the plurality of connectors defining a passage axis; and
wherein the passage axis of each of the plurality of connectors is arranged at a corresponding passage angle relative to the mean camber line, the passage angle being between 15 degrees and 90 degrees.

### Claim 13 (depends on 12)

The casting core for an airfoil as recited in claim 12, wherein the mean camber line intersects at least one of the plurality of crossover passages.

### Claim 14 (depends on 12)

The casting core for an airfoil as recited in claim 12, wherein the airfoil is a turbine airfoil.

### Claim 15

A gas turbine engine, comprising:
an array of airfoils circumferentially distributed about an engine axis, each airfoil of the array of airfoils including an airfoil section having pressure and suction sides between a leading edge and a trailing edge, the airfoil section defining a mean camber line extending between the leading and trailing edges to bisect a thickness of the airfoil section, facing pressure and suction sides of adjacent airfoils defining a throat, the throat defined as a minimum distance between the facing pressure and suction sides at a respective span position; and
wherein airfoil section comprises:
an external wall and an internal wall;
a first cavity and a second cavity separated by the internal wall, the second cavity bounded by the external wall at the leading edge;
a plurality of crossover passages within the internal wall and connecting the first cavity to the second cavity, each of the plurality of crossover passages defining a passage axis; and
wherein the passage axis of each of the plurality of crossover passages is arranged at a corresponding passage angle relative to the mean camber line such that the passage axis of each of the plurality of crossover passages intersects the external wall at a location between the throat and an aftmost location of the second cavity with respect to the engine axis.

### Claim 16 (depends on 15)

The gas turbine engine as recited in claim 15, wherein the airfoil section extends in a radial direction with respect to the engine axis from a 0% span position to a 100% span position, and an outlet of each of the plurality of crossover passage is defined at a respective span position that is between an 80% span position and the 100% span position.

### Claim 17 (depends on 16)

The gas turbine engine as recited in claim 16, comprising a plurality of film cooling passages within the external wall and extending from a surface of the second cavity, wherein the passage axis of at least some of the plurality of crossover passages intersects the external wall at an intersection point, and the second cavity is free of any film cooling passages aft of the intersection point.

### Claim 18 (depends on 17)

The gas turbine engine as recited in claim 17, wherein the passage axis of each of the plurality of crossover passages intersects the suction side.

### Claim 19 (depends on 15)

The gas turbine engine as recited in claim 15, comprising a compressor section and a turbine section, with the array of airfoils located in at least one of the compressor section and the turbine section.

### Claim 20 (depends on 19)

The gas turbine engine as recited in claim 19, wherein the array of airfoils are rotatable blades.