GAS TURBINE SYSTEM

• US 20190323432A1
• Filed: 04/05/2019
• Published: 10/24/2019
• Est. Priority Date: 04/18/2018
• Status: Active Grant

• Current Assignee: Mitsubishi Heavy Industries, Ltd.
• Original Assignee: Mitsubishi Heavy Industries, Ltd.
• Inventors: AOYAMA, Kuniaki, MASUTANI, Jo
• Granted Patent: US 11,199,135 B2
• Time in Patent Office: Days
• CPC Class Codes: 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
• CPC Class Codes: F01D 5/085 cooling fluid circulating inside the rotor
• CPC Class Codes: F01D 5/181 Blades having a closed internal cavity containing a cooling medium, e.g. sodium
• CPC Class Codes: F02C 6/06 providing compressed gas F02C6/10 takes precedence
• CPC Class Codes: F02C 7/185 Cooling means for reducing the temperature of the cooling air or gas
• CPC Class Codes: F02C 7/28 Arrangement of seals
• CPC Class Codes: F02C 9/18 by bleeding, bypassing or acting on variable working fluid interconnections between turbines or compressors or their stages F02C3/113 takes precedence
• CPC Class Codes: F05D 2240/12 Fluid guiding means, e.g. vanes
• CPC Class Codes: F05D 2260/213 by the provision of a heat exchanger within the cooling circuit

Abstract

A turbine-cooling system of a gas turbine system includes first and second intra-vane flow passages formed in stator vanes so as to penetrate them in a radial direction, an intra-rotation-shaft flow passage connecting the first intra-vane flow passage and the second intra-vane flow passage in a rotation shaft, an extra-turbine flow passage connecting the first intra-vane flow passage and the second intra-vane flow passage, a boost compressor configured to make cooling air flow sequentially through the first intra-vane flow passage, the intra-rotation-shaft flow passage, the second intra-vane flow passage, and the extra-turbine flow passage, and a cooling unit configured to cool the cooling air.

Claims

1. A gas turbine system, comprising:

a turbine including

a rotation shaft configured to rotate around an axis, a rotor blade stage including a plurality of rotor blades which are arranged on an outer circumference of the rotation shaft with intervals in a circumferential direction thereof,

a casing surrounding the rotation shaft and the plurality of rotor blades, and

a vane stage including a plurality of stator vanes which are fixed to the casing and are arranged with intervals in the circumferential direction; and

a turbine-cooling system configured to cool the turbine with cooling air,

wherein the turbine-cooling system includes

a first intra-vane flow passage that is formed in the stator vane so as to penetrate the stator vane in a radial direction of the rotation shaft,

a second intra-vane flow passage that is formed in the stator vane so as to penetrate the stator vane in the radial direction,

an intra-rotation-shaft flow passage that is formed in the rotation shaft and connects inner ends in the radial direction of the first intra-vane flow passage and the second intra-vane flow passage, and

an extra-turbine flow passage connecting an outer end in the radial direction of the first intra-vane flow passage and an outer end in the radial direction of the second intra-vane flow passage.

2. The gas turbine system according to claim 1, wherein the

turbine-cooling system further includes:

a boost compressor configured to make the cooling air flow sequentially through the first intra vane flow passage, the intra-rotation-shaft flow passage, the second intra-vane flow passage, and the extra-turbine flow passage; and

a cooling unit that is provided in the extra-turbine flow passage and configured to cool the cooling air,

3. The gas turbine system according to claim 2, wherein the turbine includes a plurality of the stator vane stages that are arranged in a main flow direction parallel to the axis, and

wherein the first intra-vane flow passage is formed in one of the plurality of the stator vane stages, and the second intra-vane flow passage is formed in another of the plurality of the stator vane stages.

4. The gas turbine system according to claim 3, wherein the first intra-vane flow passage is formed in the stator vane stage downstream in the main flow direction from the stator vane stage in which the second intra-vane flow passage is formed, and

wherein the boost compressor is installed between the intra-rotation-shaft flow passage and the second intra-vane flow passage.

5. The gas turbine system according to claim 4, wherein the boost compressor includes:

a disk-shaped impeller disk that is provided integrally with the rotation shaft and is centered on the axis;

a plurality of blades that are arranged radially centered on the axis on a surface of the impeller disk facing the axis direction; and

an impeller cover covering the blades thereby forming an impeller flow passage between the impeller disk and the impeller cover, and

wherein an inlet of the impeller flow passage is open to an outlet of the intra-rotation-shaft flow passage and an outlet of the impeller flow passage is open to an inlet of the second intra-vane flow passage.

6. The gas turbine system according to claim 2, wherein the turbine includes a plurality of the stator vane stages that are arranged in the main flow direction parallel to the axis, and

wherein the first intra-vane flow passage is formed in one of the plurality of the stator vane stages, and the second intra-vane flow passage is formed in the same of the plurality of the stator vane stages.

7. The gas turbine system according to claim 6, wherein the boost compressor includes:

a disk-shaped impeller disk that is provided integrally with the rotation shaft and is centered on the axis;

a plurality of blades that are arranged radially centered on the axis on a surface of the impeller disk facing the axis direction; and

an impeller cover covering the blades thereby forming an impeller flow passage serving as the intra-rotation-shaft flow passage between the impeller disk and the impeller cover.

8. The gas turbine system according to claim 1, wherein an intra-rotor-blade flow passage is formed in the rotor blade so as to penetrate the rotor blade in the radial direction, wherein at least a part of the cooling air which has passed through the first intra-vane flow passage flow is allowed to pass through the intra-rotor-blade flow passage.

9. The gas turbine system according to claim 1, further comprising a compressor configured to rotate along with the rotation shaft thereby compressing air serving as the cooling air,

wherein the turbine-cooling system further includes a supply flow passage configured to supply air led out from the compressor to the extra-turbine flow passage.