Turbine shroud with full hoop ceramic matrix composite blade track and seal system

Patent metadata

Field	Value
Publication number	US-10287906-B2
Application number	US-201615163079-A
Country	US
Kind code	B2
Filing date	May 24, 2016
Priority date	May 24, 2016
Publication date	May 14, 2019
Grant date	May 14, 2019

Abstract

A gas turbine engine includes a blade track, a support assembly, and a seal system. The blade track is arranged around a central axis of the gas turbine engine to direct gasses through the gas turbine engine. The support assembly is arranged around the blade track to support the blade track relative to the central axis. The seal system is configured to block fluid communication between pressurized cavities formed in the support assembly.

First claim

What is claimed is:

1. A gas turbine engine comprising
a blade track arranged around a central axis of the gas turbine engine, the blade track having a leading edge, a trailing edge axially spaced apart from the leading edge, a radial outer surface that extends between the leading and trailing edges, and a radial inner surface spaced apart from the radial outer surface,
a support assembly arranged around the blade track to support the blade track relative to the central axis, the support assembly being formed to define a retention cavity configured to receive relatively high-pressure compressor air that is directed to the leading edge of the blade track to resist gasses from flowing past the leading edge over the radial outer surface of the blade track, a vent cavity configured to receive relatively low-pressure compressor air that is directed to the trailing edge of the blade track to resist the gasses from flowing past the trailing edge over the radial outer surface of the blade track, and a bleed cavity fluidly interconnecting the retention cavity with the vent cavity and configured to receive intermediate-pressure compressor air, and
a seal system including a first seal configured to block fluid communication between the retention cavity and the bleed cavity and a second seal configured to block fluid communication between the bleed cavity and the vent cavity so that a pressure difference across either of the first and second seals is less than a pressure difference between the retention cavity and the vent cavity.

2. The gas turbine engine of claim 1 , further comprising a compressor having a plurality of compressor stages and the bleed cavity is fluidly connected to a first of the plurality of compressor stages.

3. The gas turbine engine of claim 2 , wherein the retention cavity is fluidly connected to a second of the plurality of compressor stages and the second of the plurality of compressor stages is located downstream of the first of the plurality of compressor stages.

4. The gas turbine engine of claim 1 , wherein the support assembly includes a carrier arranged radially outward of and around the blade track and a track-retaining vane assembly that cooperate with the blade track to define the vent cavity, the track-retaining vane assembly including an outer band that engages the trailing edge of the blade track and a static airfoil that extends radially inward of the outer band.

5. The gas turbine engine of claim 4 , wherein the support assembly further includes a turbine case arranged around the carrier and the turbine case cooperates with the carrier to define the retention cavity.

6. The gas turbine engine of claim 4 , wherein the support assembly further includes a turbine case arranged around the track-retaining vane assembly and the carrier and the track-retaining vane assembly cooperate to define the bleed cavity.

7. The gas turbine engine of claim 1 , wherein the support assembly includes a carrier that cooperates with the blade track to define a thermal management chamber therebetween.

8. The gas turbine engine of claim 7 , wherein the support assembly further includes a distribution plate coupled to the carrier, the distribution plate extends axially through the thermal management chamber to partition the thermal management chamber into a radial outer distribution cavity configured to receive and distribute cooling air circumferentially around the central axis and a radially inner cooling cavity, and the distribution plate is formed to include a plurality of holes arranged to direct the cooling air out of the distribution cavity into the cooling cavity toward the radial outer surface of the blade track.

9. The gas turbine engine of claim 8 , wherein cooling cavity is in fluid communication with the vent cavity.

10. The gas turbine engine of claim 8 , further comprising a compressor having a plurality of compressor stages, the thermal management chamber is fluidly connected to a first of the plurality of compressor stages, the bleed cavity is fluidly connected to a second of the plurality of compressor stages, the retention cavity is fluidly connected to a third of the plurality of compressor stages, the third of the plurality of compressor stages is located downstream of the first and second of the plurality of compressor stages, and the second of the plurality of compressor stages is located downstream of the first of the plurality of compressor stages.

11. A gas turbine engine comprising
a turbine case arranged around a central axis of the gas turbine engine,
an annular turbine shroud arranged to define an inner flow path configured to conduct hot, low-pressure gasses axially through the annular turbine shroud, the annular turbine shroud located radially between the turbine case and the central axis to define a retention cavity located between the annular turbine shroud and the turbine case,
a blade-track retainer located radially between the turbine case and the central axis to define a bleed cavity located between the blade-track retainer and the turbine case and the blade-track retainer and the annular turbine shroud cooperate to define a vent cavity in fluid communication with the low-pressure gasses in the inner flow path,
a first seal engaged with the turbine case and the turbine shroud to block air from escaping the retention cavity and flowing into one of the bleed cavity and the vent cavity, and
a second seal engaged with the turbine shroud and the blade-track retainer to block air from escaping the bleed cavity and flowing into the vent cavity so that a pressure difference across either of the first and second seals is less than a pressure difference between a pressure of the air in the retention cavity and a pressure of the low-pressure gasses.

12. The gas turbine engine of claim 11 , wherein the annular turbine shroud includes a full hoop blade track comprising ceramic materials and a carrier comprising metallic materials and arranged around the blade track.

13. The gas turbine engine of claim 12 , wherein the carrier includes a support band that extends axially relative to the central axis to form a portion of the retention cavity and an aft flange that extends radially outward toward the turbine case to close an aft end of the retention cavity and the first seal engages the aft flange of the carrier.

14. The gas turbine engine of claim 13 , wherein the turbine case includes an outer band spaced apart from the support band of the carrier and a partition flange that extends radially inward toward the central axis, the partition flange is located adjacent the aft flange of the carrier, and the first seal further engages the partition flange.

15. The gas turbine engine of claim 13 , wherein the blade-track retainer includes a forward flange that extends radially outward toward the turbine case to close a forward end of the bleed cavity and the second seal engages the aft flange of the carrier and the forward flange of the blade-track retainer.

16. The gas turbine engine of claim 11 , further comprising a compressor having a plurality of compressor stages and the bleed cavity is fluidly connected to a first of the plurality of compressor stages.

17. The gas turbine engine of claim 16 , wherein the retention cavity is fluidly connected to a second of the plurality of compressor stages and the second of the plurality of compressor stages is located downstream of the first of the plurality of compressor stages.

18. The gas turbine engine of claim 11 , wherein the annular turbine shroud includes a blade track, a carrier arranged circumferen

Assignees

• Rolls Royce Nam Tech Inc

Inventors

• Rice Edward C
• Vetters Daniel K

Classifications

• F05D2220/32

  in gas turbines · CPC title

• F01D11/005 Primary

  Sealing means between non relatively rotating elements · CPC title

• F01D25/24

  Casings (modified for heating or cooling F01D25/14); Casing parts, e.g. diaphragms, casing fastenings (casings for rotary machines or engines in general F16M {; special arrangements in stators dealing with breaking-off of part of rotor F01D21/045}) · CPC title

• F05D2240/11

  Shroud seal segments · CPC title

• F05D2300/10

  Metals, alloys or intermetallic compounds · CPC title