Passive blade tip clearance control system for gas turbine engine

Patent metadata

Field	Value
Publication number	US-2020208533-A1
Application number	US-201816233964-A
Country	US
Kind code	A1
Filing date	Dec 27, 2018
Priority date	Dec 27, 2018
Publication date	Jul 2, 2020
Grant date	—

Abstract

The present disclosure relates to a gas turbine engine including a turbine wheel mounted for rotation about a central axis and a turbine shroud ring mounted radially outward from the turbine wheel. The turbine wheel includes a plurality of blades that are spaced apart radially from the turbine shroud ring to establish a blade tip clearance gap. The gas turbine engine further includes a blade tip clearance control system that passively controls the size of the clearance gap based on engine operation.

First claim

What is claimed is:

1 . A gas turbine engine comprising
a compressor configured to pressurize air moving along a primary gas path of the gas turbine engine, a combustor fluidly coupled to the compressor to receive pressurized air discharged from the compressor and configured to ignite fuel mixed with the pressurized air, and a turbine including (i) a high-pressure section fluidly coupled to the combustor to receive combustion gases generated by fuel burned in the combustor and (ii) a low-pressure section fluidly coupled to receive the combustion gasses exiting the high-pressure section, wherein the high-pressure section includes a turbine wheel mounted for rotation about a central reference axis, a variable-diameter turbine shroud ring that extends around the turbine wheel, and a passive blade-tip clearance control system including a shroud-ring support coupled to the variable-diameter turbine shroud ring that is configured to drive motion of the turbine shroud ring radially inward or outward based on temperature of the shroud-ring support and defining at least in part a cavity located radially outward of the variable-diameter turbine shroud ring, and wherein the cavity is fluidly coupled to a bleed-air passageway that extends from the compressor to the cavity without interruption from a valve and a a cooling-air passageway that extends from the cavity to the low pressure section such that pressurized bleed air from the compressor is conducted to the cavity of the passive blade tip clearance control system so that the temperature and motion of the shroud-ring support is controlled based on the operating conditions of the engine without active control of the pressurized bleed air provided to the cavity.

2 . The gas turbine engine of claim 1 , wherein the passive blade tip clearance control system further includes an outer case and the shroud-ring support is provided by an inner case mounted radially-inward of the outer case to define the cavity radially therebetween.

3 . The gas turbine engine of claim 2 , wherein the passive blade tip clearance control system further includes an inlet conduit coupled to the outer case and opening into the cavity and an outlet, the inlet configured to conduct the bleed air from the compressor into the cavity and the outlet configured to conduct the bleed air from the cavity to the low pressure section of the turbine.

4 . The gas turbine engine of claim 3 , wherein the passive blade-tip clearance control system is configured to heat the inner case during start-up conditions of the gas turbine engine and is configured to cool the inner case during cruise conditions.

5 . The gas turbine engine of claim 2 , wherein the cavity formed between the outer case and the inner case is sealed off from a gas path of the high pressure section of the turbine such that the temperature of gases within the cavity controls the gap while allowing for pressure within the cavity to be less than pressure within the primary gas path of the high pressure turbine section.

6 . The gas turbine engine of claim 2 , wherein the outer case includes an annular duct that extends circumferentially around the reference axis and defines a manifold and the inlet is fluidly coupled to the manifold.

7 . The gas turbine engine of claim 6 , wherein the passive blade tip clearance control system includes a plurality of inlet conduits fluidly coupled to the manifold and spaced apart circumferentially around the reference axis and a plurality of outlets spaced apart circumferentially around the reference axis that extend through the manifold and are offset from each inlet conduit.

8 . The gas turbine engine of claim 2 , wherein the high pressure section of the turbine includes a first turbine blade stage, a second turbine blade stage axially aft of the first turbine blade stage, and a vane stage axially between the first and second turbine blade stages, and the passive blade tip clearance control system is configured to control a gap radially between second turbine blade stage and the turbine shroud ring.

9 . The gas turbine engine of claim 8 , wherein the outer case includes an outer panel spaced apart from the central reference axis a first distance, and an inner panel spaced apart from the central reference axis a second distance that is less than the first distance.

10 . The gas turbine engine of claim 9 , wherein the inner panel is positioned radially outward of the second turbine blade stage such that the cavity is narrowed outward of the second turbine blade stage.

11 . The gas turbine engine of claim 9 , wherein the outer panel is spaced apart axially from the inner panel.

12 . The gas turbine engine of claim 9 , wherein the inner panel is adjustable axially to target additional turbine blade stages included in the high pressure section of the turbine.

13 . The gas turbine engine of claim 8 , wherein the inner case includes a plurality of turbulators coupled to an upper surface of the inner case within the cavity radially outward of the second turbine blade stage to increase heat transfer between the bleed air and the inner case directly outward of the second turbine blade stage.

14 . The gas turbine engine of claim 2 , wherein the inner case includes a panel that is coupled to the turbine shroud ring and a flange coupled to an axially-forward end of the panel, the flange coupled to the outer case and having a U-shape when viewed circumferentially so that the flange is configured to flex as the inner case moves radially inward and outward relative to the outer case.

15 . A high pressure turbine section for use in a gas turbine engine, the turbine section comprising
a turbine wheel mounted for rotation about a central reference axis, a plurality of blades that extend radially outward from the turbine wheel to interact with gases moving through a primary gas path of the turbine section, a variable-diameter turbine shroud ring that extends around the turbine wheel to define a radially-outer boundary of the primary gas path, and a passive blade-tip clearance control system configured to drive motion of the turbine shroud ring radially inward and outward relative to the central reference axis to control size of a gap radially between the turbine wheel and the variable-diameter turbine shroud ring, the passive blade-tip clearance control system including an outer case and an inner case mounted radially-inward of the outer case to define a cavity radially therebetween, wherein the cavity formed between the outer case and the inner case is sealed off from the primary gas path within the high pressure turbine section.

16 . The gas turbine engine of claim 15 , wherein the outer case includes an outer panel spaced apart from the central reference axis a first distance, and an inner panel spaced apart from the central reference axis a second distance that is less than the first distance.

17 . The gas turbine engine of claim 16 , wherein the high pressure section of the turbine includes a first turbine blade stage, a second turbine blade stage axially aft of the first turbine blade stage, and a vane stage axially between the first and second turbine blade stages, and wherein the inner panel is positioned radially outward of the second turbine blade stage such that the cavity is narrowed outboard of the second turbine blade stage.

18 . The gas turbine engine of claim 16 , wherein the inner panel is adjustable axially to target additional turbine blade stages included in the high pressure section of the turbine.

19 . The gas tur

Assignees

• Rolls Royce Corp
• Rolls Royce Nam Tech Inc

Inventors

• Kapadia Behram
• Moody Jack
• Humes Ryan C
• Snyder Brandon R

Classifications

• F02C9/18

  by bleeding, bypassing or acting on variable working fluid interconnections between turbines or compressors or their stages {(F02C3/113 takes precedence)} · CPC title

• F02K3/06

  with front fan · CPC title

• F01D11/18 Primary

  using stator or rotor components with predetermined thermal response, e.g. selective insulation, thermal inertia, differential expansion · CPC title

• F05D2220/323

  for aircraft propulsion, e.g. jet engines · CPC title

• F05D2260/606

  Bypassing the fluid · CPC title