Full hoop blade track with interstage cooling air

Patent metadata

Field	Value
Publication number	US-10240476-B2
Application number	US-201615000661-A
Country	US
Kind code	B2
Filing date	Jan 19, 2016
Priority date	Jan 19, 2016
Publication date	Mar 26, 2019
Grant date	Mar 26, 2019

Abstract

A gas turbine engine includes a turbine having a plurality of vanes, a plurality of blades, a turbine shroud arranged around the vanes and blades, and a turbine case arranged around the turbine shroud. The turbine shroud is sized to block combustion products from passing over the blades without pushing the blades to rotate. The turbine shroud includes a runner arranged around the blades and a carrier arranged around the runner.

First claim

What is claimed is:

1. A turbine shroud for use in a gas turbine engine having a central axis, the turbine shroud comprising
an annular carrier formed to define a radially inwardly-opening carrier channel that extends around the central axis and the annular carrier includes an outer pin receiver that extends through the annular carrier and opens into the carrier channel to allow pressurized cooling air to pass through the annular carrier into the carrier channel and a high-pressure cooling air passageway that extends radially through the annular carrier,
a one-piece annular runner aligned axially with the carrier channel of the annular carrier, the one-piece annular runner includes an inner radial runner surface located radially between the annular carrier and the central axis and an outer radial runner surface located radially between the inner radial runner surface and the annular carrier, and the outer radial runner surface cooperates with the annular carrier to form an annular buffer chamber between the annular carrier and the one-piece annular runner, and
a cooling system including an annular impingement plate positioned in the annular buffer chamber to separate the annular buffer chamber into an outer chamber and an inner chamber located radially between the outer chamber and the one-piece annular runner, the outer pin receiver opens into the outer chamber to direct the pressurized cooling air into the outer chamber, and the annular impingement plate includes a plurality of diffusion holes spaced circumferentially around the annular impingement plate and each diffusion hole extends radially through the impingement plate to direct the pressurized cooling air in the outer chamber through the annular impingement plate into the inner chamber and toward the outer radial runner surface of the one-piece annular runner,
wherein the one-piece annular runner includes a forward section, an aft section spaced apart axially from the forward section, and a midsection extending between the forward section and the aft section, the high-pressure cooling air passageway is configured to direct high-pressure air toward the forward section of the one-piece annular runner, the high-pressure air has a greater pressure than the pressurized cooling air, and the turbine shroud further includes a first seal positioned radially between the one-piece annular runner and the annular carrier and positioned axially between the high-pressure cooling air passageway and the annular buffer chamber.

2. The turbine shroud of claim 1 , wherein the one-piece annular runner includes circumferentially spaced apart hot zones associated with relatively high temperatures during operation of the gas turbine engine and each diffusion hole formed in the annular impingement plate is arranged to direct the pressurized cooling air toward a corresponding hot zone.

3. The turbine shroud of claim 1 , wherein the diffusion holes formed in the annular impingement plate are arranged to direct the pressurized cooling air toward at least one of the aft section and the midsection.

4. The turbine shroud of claim 1 , wherein the gas turbine engine includes a turbine case arranged around the annular carrier, the cooling system further includes a hollow insert pin configured to extend through the turbine case and the outer pin receiver formed in the annular carrier to couple the annular carrier to the turbine case, and the hollow insert pin is configured to direct the pressurized cooling air through the turbine case and the annular carrier into the outer chamber.

5. The turbine shroud of claim 4 , wherein the cooling system further includes a controller configured to modulate a flow rate of the pressurized cooling air directed through the hollow insert pin into the outer chamber to control an expansion and contraction of the one-piece annular runner.

6. The turbine shroud of claim 4 , further including a second seal positioned between the outer radial runner surface of the one-piece annular runner and the annular carrier to block the pressurized cooling air from escaping the inner chamber and the second seal is positioned axially aft of the first seal to locate the annular buffer chamber axially between the first seal and the second seal.

7. The turbine shroud of claim 6 , wherein the second seal includes a piston ring made from one of a metallic material, a ceramic material, and a ceramic matrix composite material.

8. A gas turbine engine comprising
a turbine case arranged around a central axis of the gas turbine engine, the turbine case includes one or more outer keyways that extend through the turbine case,
a turbine shroud including (i) a carrier formed to define a radially inwardly-opening carrier channel that extends around the central axis and one or more outer pin receivers that extend through the carrier and open into the carrier channel and (ii) an annular runner aligned axially with the carrier and positioned to close the inwardly-opening carrier channel to form an annular buffer chamber between the carrier and the annular runner, and
a cooling system including one or more hollow outer insert pins that extend through the corresponding one or more outer keyways formed in the turbine case and the one or more outer pin receivers formed in the carrier into the buffer chamber to allow pressurized cooling air to pass through the turbine case and the carrier into the buffer chamber,
wherein the annular runner includes a forward section, an aft section spaced apart axially from the forward section, and a midsection extending between the forward section and the aft section, the carrier is formed to include a high-pressure cooling air passageway that extends through the carrier and is configured to direct high-pressure air toward the forward section of the annular runner, and the high-pressure air has a greater pressure than the pressurized cooling air,
wherein the gas turbine engine further includes a first seal positioned radially between the annular runner and the carrier and positioned axially between the high-pressure cooling air passageway and the buffer chamber.

9. The gas turbine engine of claim 8 , further including a second seal positioned radially between the annular runner and the carrier and positioned axially aft of the first seal to locate the buffer chamber axially between the first and the second seal and each of the first seal and the second seal includes a piston ring made from one of a ceramic and a ceramic matrix composite material.

10. The gas turbine engine of claim 8 , wherein the cooling system further includes an impingement plate positioned in the annular buffer chamber to separate the annular buffer chamber into an outer chamber and an inner chamber located radially between the outer chamber and the annular runner, the one or more hollow outer insert pins is in fluid communication with the outer chamber to direct the pressurized cooling air into the outer chamber, and the impingement plate includes a plurality of diffusion holes spaced apart circumferentially around the impingement plate that extend radially through the impingement plate to direct the pressurized cooling air in the outer chamber through the impingement plate into the inner chamber and toward the annular runner.

11. The gas turbine engine of claim 10 , wherein the annular runner includes circumferentially spaced apart hot zones associated with relatively high temperatures during operation of the gas turbine engine and each diffusion hole is formed to direct the pressurized cooling air toward a corresponding hot zone.

12. The gas turbine engine of claim 10 , wherein the plurality of diffusion holes are arranged to direct the pressurized cooling air towar

Assignees

• Rolls Royce Corp
• Rolls Royce Nam Tech Inc

Inventors

• Varney Bruce E
• Petty Jack D
• Vetters Daniel K
• Clemens Eugene

Classifications

• Y02T50/672

  Cross-Sectional Technologies · mapped topic

• F01D11/24 Primary

  by selectively cooling-heating stator or rotor components · CPC title

• F01D25/246

  Fastening of diaphragms or stator-rings · CPC title

• F05D2300/6033

  Ceramic matrix composites [CMC] · CPC title

• F01D11/005

  Sealing means between non relatively rotating elements · CPC title