Tip clearance control for turbine blades

• US 10,641,122 B2
• Filed: 08/06/2018
• Issued: 05/05/2020
• Est. Priority Date: 12/16/2014
• Status: Active Grant

• Current Assignee: Rolls-Royce PLC (Rolls-Royce Holdings Plc)
• Original Assignee: Rolls-Royce PLC (Rolls-Royce Holdings Plc)
• Inventors: Lewis, Leo V.
• Primary Examiner(s): Edgar, Richard A
• Assistant Examiner(s): Adjagbe, Maxime M
• Application Number:
  US16/055,466
• Publication Number:
  US 20180340442A1
• Time in Patent Office: 638 Days
• CPC Class Codes: F01D 11/24 by selectively cooling-heating stator or rotor components
• CPC Class Codes: F01D 25/10 Heating, e.g. warming-up before starting
• CPC Class Codes: F01D 25/12 Cooling
• CPC Class Codes: F01D 25/24 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 Class Codes: F01D 5/02 Blade-carrying members, e.g. rotors rotors of non-bladed type F01D1/34; stators F01D9/00 ; selecting particular materials F01D5/28
• CPC Class Codes: F01D 5/12 Blades specially adapted for radial flow machines or engines F01D5/04; blade roots F01D5/30; rotors with blades adjustable in operation F01D7/00; stator blades F01D9/02
• CPC Class Codes: F05D 2220/32 in gas turbines
• CPC Class Codes: F05D 2260/201 by impingement of a fluid

Abstract

An arrangement for a gas turbine engine that includes a turbine blade configured to rotate about an axis, a casing radially outside of the turbine blade, and a carrier segment mounted to the casing so as to define a first impingement space therebetween. The carrier segment is positioned radially outside the turbine blade and includes a first impingement carrier wall, a main carrier wall, and a cooling chamber. The first impingement carrier wall is adjacent to and radially inside of the first impingement space, and the first impingement carrier wall includes a first aperture. The main carrier wall is radially inside of the first impingement carrier wall. The cooling chamber is radially inside of the main carrier wall. Additionally, an intermediate chamber is disposed radially between the cooling chamber and the first impingement space. A second aperture is configured to allow ingress of air into the intermediate chamber.

Claims

1. An arrangement for a gas turbine engine comprising:

a. a turbine blade configured to rotate about an axis;

b. a casing radially outside the turbine blade;

c. a carrier segment mounted to the casing so as to define a first impingement space therebetween, the carrier segment being positioned radially outside the turbine blade and comprising;

i. a first impingement carrier wall adjacent to and radially inside of the first impingement space, the first impingement carrier wall including a first aperture;

ii. a main carrier wall radially inside of the first impingement carrier wall;

iii. a cooling chamber radially inside of the main carrier wall;

iv. an intermediate chamber radially between the cooling chamber and the first impingement space;

v. a second aperture configured to allow ingress of air into the intermediate chamber, and wherein the first aperture of the first impingement carrier wall is configured to allow ingress of the air into the first impingement space; and

vi. a third aperture configured to allow flow of the air from the first impingement space to the cooling chamber.

2. The arrangement of claim 1 wherein the intermediate chamber is tapered in an axial direction of the gas turbine engine.

3. The arrangement of claim 2 wherein the intermediate chamber is larger proximal to the second aperture than distal to the second aperture.

4. The arrangement of claim 1, wherein the third aperture is formed in an axially downstream end portion of the carrier segment.

5. The arrangement of claim 1, further comprising a second impingement space that is defined between the casing and a second impingement carrier wall spaced radially inwardly from the casing and axially adjacent to the first impingement carrier wall, the second impingement carrier wall comprising a plurality of impingement apertures configured to allow ingress of air into the second impingement space.

6. The arrangement of claim 5, further comprising a fourth aperture that allows air flow between the second impingement space and the first impingement space.

7. The arrangement of claim 1 further comprising a fifth aperture in the main carrier wall that is configured to allow air flow between the intermediate chamber and the cooling chamber.

8. A gas turbine engine comprising the arrangement of claim 1.

9. A method of controlling a temperature of a turbine casing of a gas turbine engine,

the engine including:

a turbine blade configured to rotate about an axis;

a casing radially outside the turbine blade;

a carrier segment mounted to the casing so as to define a first impingement space therebetween, the carrier segment being positioned radially outside the turbine blade and including;

a first impingement carrier wall adjacent to and radially inside of the first impingement space, the first impingement carrier wall including a first aperture,

a main carrier wall radially inside of the first impingement carrier wall;

a cooling chamber radially inside of the main carrier wall; and

an intermediate chamber radially between the cooling chamber and the first impingement space;

the method comprising;

passing air from an air feed source to the intermediate chamber;

passing the air from the intermediate chamber, through the first aperture in the first impingement carrier wall, and into the first impingement space so that the air impinges on the casing; and

passing the air from the first impingement space into the cooling chamber.

10. The method of claim 9, wherein a second aperture is disposed between the air feed source and the intermediate chamber, and the air passes through the second aperture to reach the intermediate chamber.

11. The method of claim 9, wherein a third aperture is disposed between the first impingement space and the cooling chamber such that the air passes directly from the first impingement space, through the third aperture, and into the cooling chamber.

12. The method of claim 9, further comprising passing air directly from the intermediate chamber to the cooling chamber.