# US 10,502,093 B2 · Turbine shroud cooling

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- Number: US 10,502,093 B2
- Document ID: US10502093B2
- Status: Granted
- Kind: B2
- Application number: 15/840,492
- Filed: 2017-12-13
- Published: 2019-06-13
- Granted: 2019-12-10
- Earliest claimed: 2017-12-13
- Assignees: Pratt & Whitney Canada Corp.
- Applicants: Pratt & Whitney Canada Corp.
- Inventors: Remy Synnott; Mohammed Ennacer; Chris Pater; Denis Blouin; Kapila Jain; Farough Mohammadi
- Web page: https://patentgrep.com/patent/US10502093B2

## CPC

- F01D 25/12: Cooling
- B22C 9/10: Cores; Manufacture or installation of cores {(breaker cores B22C9/084)}
- B22C 9/24: for hollow articles
- F01D 5/081: {Cooling fluid being directed on the side of the rotor disc or at the roots of the blades (F01D5/087 takes precedence)}
- F01D 5/185: {Liquid cooling (F01D5/181 takes precedence)}
- F01D 5/225: {by shrouding}
- F01D 25/246: {Fastening of diaphragms or stator-rings}
- F05D 2230/211: by precision casting, e.g. microfusing or investment casting
- F05D 2240/11: Shroud seal segments
- F05D 2260/201: by impingement of a fluid
- F05D 2260/202: by film cooling
- F05D 2260/205: Cooling fluid recirculation, i.e. after cooling one or more components is the cooling fluid recovered and used elsewhere for other purposes
- F05D 2260/2212: by creating turbulence
- F05D 2260/22141: using fins or ribs
- F01D: NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES (machines or engines for liquids F03; non-positive displacement pumps F04D)
- B22C: FOUNDRY MOULDING (moulding refractory materials in general B28B)
- F05D: INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS

## Abstract

A turbine shroud segment has a body extending axially between a leading edge and a trailing edge and circumferentially between a first and a second lateral edge. A core cavity is defined in the body and extends axially from a front end adjacent the leading edge to a rear end adjacent to the trailing edge. A plurality of cooling inlets and outlets are respectively provided along the front end and the rear end of the core cavity. A crossover wall extends across the core cavity and defines a row of crossover holes configured to accelerate the flow of coolant directed into the core cavity via the cooling inlets. The crossover wall is positioned to accelerate the coolant flow at the beginning of the cooling scheme where the shroud segment is the most thermally solicited.

## Claims

### Claim 1

A turbine shroud segment for a gas turbine engine having an annular gas path extending about an engine axis, the turbine shroud segment comprising: a body extending axially between a leading edge and a trailing edge and circumferentially between a first and a second lateral edge; a core cavity defined in the body and extending axially from a front end adjacent the leading edge to a rear end adjacent to the trailing edge; a plurality of cooling inlets along the front end of the core cavity; a plurality of cooling outlets along the rear end of the core cavity; and a crossover wall extending across the core cavity and defining a row of crossover holes forming a constriction to accelerate a flow of coolant delivered into the core cavity by the cooling inlets, the crossover wall being positioned axially closer to the cooling inlets than the cooling outlets.

### Claim 2 (depends on 1)

The turbine shroud segment defined in claim 1, wherein the row of crossover holes comprises two distinct sets of crossover holes, a first set including laterally outermost holes positioned at a boundary of the core cavity along the first and second lateral edges of the body, and a second set including intermediate holes positioned between the laterally outermost holes, the laterally outermost holes being configured to direct the coolant passing therethrough onto an interior side of the first and second lateral edges, the intermediate holes being configured to direct the coolant in an area of the core cavity intermediate between the first and second lateral edges of the body.

### Claim 3 (depends on 2)

The turbine shroud segment defined in claim 2, wherein the laterally outermost holes and the intermediate holes have a different cross-sectional area.

### Claim 4 (depends on 3)

The turbine shroud segment defined in claim 3, wherein the laterally outermost holes have a greater cross-sectional area than that of the intermediate holes.

### Claim 5 (depends on 4)

The turbine shroud segment defined in claim 4, wherein the laterally outermost holes extend along the interior side of the first and second lateral edges and have a different cross-sectional shape than that of the intermediate holes.

### Claim 6 (depends on 2)

The turbine shroud segment defined in claim 2, wherein the laterally outermost holes are impingement holes configured to cause coolant to impinge upon the interior side of the first and second lateral edges of the body.

### Claim 7 (depends on 2)

The turbine shroud segment defined in claim 2, wherein the laterally outermost holes are angled with respect to the first and second lateral edges and define a feed direction aiming at a hottest area along the first and second lateral edges of the body.

### Claim 8 (depends on 2)

The turbine shroud segment defined in claim 2, wherein the laterally outermost holes have an oblong cross-section, and wherein the intermediate holes have a circular cross-section.

### Claim 9 (depends on 1)

The turbine shroud segment defined in claim 1, wherein the crossover holes have a smaller cross-sectional area than that of the plurality of cooling inlets.

### Claim 10 (depends on 1)

The turbine shroud segment defined in claim 1, further comprising turning vanes in opposed corners of the front end of the core cavity.

### Claim 11 (depends on 10)

The turbine shroud segment defined in claim 10, wherein the turning vanes are positioned upstream of the crossover wall relative to the flow of coolant though the core cavity.

### Claim 12 (depends on 11)

The turbine shroud segment defined in claim 11, wherein the plurality of cooling inlets are inclined so as to define a feed direction having an axial component pointing in an upstream direction relative to the flow of coolant through the core cavity.

### Claim 13 (depends on 1)

The turbine shroud segment defined in claim 1, further comprising a plurality of pedestals extending integrally from a bottom wall of the core cavity to a top wall thereof, the bottom wall corresponding to a back side of a radially inner wall of the body, the top wall corresponding to the back side of a radially outer wall of the body, the body being monolithic.

### Claim 14 (depends on 13)

The turbine shroud segment defined in claim 13, wherein the plurality of pedestals includes a first set of pedestals positioned upstream of the crossover wall and a second set of pedestals positioned downstream of the crossover walls.

### Claim 15

A method of manufacturing a turbine shroud segment comprising: using a casting core to create an internal cooling circuit of the turbine shroud segment, the casting core having a body including a front portion connected to a rear portion by a transverse row of pins, the transverse row of pins including lateral pins positioned along opposed lateral edges of the body, the lateral pins having a greater cross-sectional area than that of the other pins of the transverse row of pins, and a plurality of holes defined through the front portion and the rear portion of the body of the casting core; casting a body of the turbine shroud segment about the casting core; and removing the casting core from the cast body of the turbine shroud segment.

### Claim 16 (depends on 15)

The method defined in claim 15, wherein the casting core further comprises a transverse row of ribs extending from a top surface of the front portion of the body of the casting core, and wherein the method comprises using the casting core to form as-cast inlet passages in a front portion of the turbine shroud segment.

### Claim 17 (depends on 15)

The method defined in claim 15, wherein the casting core further comprises a transverse row of pins projecting from a rear end of the rear portion of the body of the casting core, and wherein the method comprises using the casting core to form as-cast outlet passages in a trailing edge of the turbine shroud segment.