Gas turbine system

Patent metadata

Field	Value
Publication number	US-11199135-B2
Application number	US-201916376054-A
Country	US
Kind code	B2
Filing date	Apr 5, 2019
Priority date	Apr 18, 2018
Publication date	Dec 14, 2021
Grant date	Dec 14, 2021

Abstract

A turbine-cooling system of a gas turbine system includes a first intra-vane flow passage defined in a first stator vane so as to penetrate the first stator vane in a radial direction, a second intra-vane flow passage defined in a second stator vane so as to penetrate the second stator vane in the 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.

First claim

What is claimed is:

1. A gas turbine system, comprising:
a turbine including
a rotation shaft configured to rotate around an axis defining a circumferential direction and a radial direction,
a rotor blade stage including a plurality of rotor blades arranged on an outer circumference of the rotation shaft with intervals in the circumferential direction,
a casing surrounding the rotation shaft and the plurality of rotor blades, and
a plurality of stator vane stages, each including a plurality of stator vanes fixed to the casing and 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 defined in a first vane of a first vane stage of the plurality of stator vane stages so as to penetrate the first vane in the radial direction,
a second intra-vane flow passage that is defined in a second vane of a second vane stage of the plurality of stator vane stages so as to penetrate the second vane in the radial direction,
an intra-rotation-shaft flow passage that is defined in the rotation shaft and connects an inner end in the radial direction of the first intra-vane flow passage and an inner end in the radial direction of the second intra-vane flow passage,
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, and
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,

wherein the boost compressor includes an impeller disk that is centered on the axis and integral with the rotation shaft such that the impeller disk protrudes directly radially outward from the rotation shaft,
wherein an inlet of the boost compressor faces downstream in a main flow direction of the turbine, and
wherein the boost compressor is between the intra-rotation-shaft flow passage and the second intra-vane flow passage.

2. The gas turbine system according to claim 1 , wherein the turbine-cooling system further includes:
a cooling unit that is in the extra-turbine flow passage and configured to cool the cooling air.

3. The gas turbine system according to claim 1 , wherein the boost compressor further includes:
a plurality of blades radially centered on the axis on a surface of the impeller disk facing downstream in the main flow direction; and
an impeller cover covering the plurality of blades and defining an impeller flow passage between the impeller disk and the impeller cover,
wherein the inlet of the boost compressor which is 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.

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

5. The gas turbine system according to claim 1 , further comprising a compressor configured to rotate along with the rotation shaft so as to compress air,
wherein the turbine-cooling system further includes a supply flow passage configured to supply the air from the compressor to the extra-turbine flow passage.

Assignees

• Mitsubishi Heavy Ind Ltd

Inventors

• Aoyama Kuniaki
• Masutani Jo

Classifications

• F02C7/28

  Arrangement of seals · CPC title

• 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

• F01D5/085

  cooling fluid circulating inside the rotor · CPC title

• F05D2240/12

  Fluid guiding means, e.g. vanes · CPC title

• F02C7/185 Primary

  Cooling means for reducing the temperature of the cooling air or gas · CPC title