Gas turbine stationary blade and gas turbine

A turbine stationary blade, comprising:

• a blade body;
• and
• a shroud formed in an end portion of the blade body in a blade height direction,
• wherein the shroud includes:
• a bottom plate in contact with a combustion gas flow passage;
• a peripheral wall formed in the blade height direction and along a peripheral edge of the bottom plate;
• a recess forming a space surrounded by the peripheral wall and the bottom plate;
• a plurality of partition ribs connecting the blade body and the peripheral wall, dividing the recess into a plurality of spaces, and forming a plurality of cavities;
• and
• an impingement plate dividing the space into an outer cavity formed on an outer side in the blade height direction and an inner cavity formed on an inner side of the outer cavity, and having a plurality of through holes via which the outer cavity and the inner cavity communicate with each other,
• wherein the peripheral wall includes:
• a leading edge end portion extending to a leading edge side of the blade body;
• a suction side end portion extending from a leading edge to a trailing edge on a suction side of the blade body, and internally including a suction side passage;
• a pressure side end portion extending from a leading edge to a trailing edge on a pressure side of the blade body, and internally including a pressure side passage;
• and
• a trailing edge end portion extending to a trailing edge side of the blade body,
• wherein the plurality of cavities are formed from the inner cavity, and include:
• a first cavity disposed on a leading edge pressure side so as to surround a circumference of the blade body between the blade body and the peripheral wall, connected to a cooling hole formed in the bottom plate, and connected to the pressure side passage;
• and
• a fourth cavity disposed adjacent to an axially downstream side of the first cavity, disposed on a most axially upstream side of the pressure side among the partition ribs, disposed via a first partition rib connecting the pressure side end portion and the pressure surface of the blade body, and having a plurality of cooling holes in the bottom plate,
• wherein an opening density of the through holes of the impingement plate in the fourth cavity is lower than the opening density of the through holes of the impingement plate in the first cavity,
• wherein the fourth cavity includes a plurality of intermediate cavities which are divided by a second partition rib constituted by a plurality of intermediate partition ribs connecting the pressure side end portion and the pressure surface of the blade body, and are disposed along the pressure surface of the blade body from an axially upstream side toward the axially downstream side, and
• wherein the opening density of the through holes of the impingement plate in the intermediate cavities decreases toward the axially downstream side, and is lower in the intermediate cavities disposed axially downstream via the intermediate partition ribs than in the intermediate cavities on the axially upstream side.

A cooling method for a turbine stationary blade including:

• a blade body;
• and
• a shroud formed in an end portion of the blade body in a blade height direction,
• the shroud including:
• a bottom plate in contact with a combustion gas flow passage;
• a peripheral wall formed in the blade height direction and along a peripheral edge of the bottom plate;
• a recess forming a space surrounded by the peripheral wall and the bottom plate;
• a plurality of partition ribs connecting the blade body and the peripheral wall, dividing the recess into a plurality of spaces, and forming a plurality of cavities;
• and
• an impingement plate dividing the space into an outer cavity formed on an outer side in the blade height direction and an inner cavity formed on an inner side of the outer cavity, and having a plurality of through holes via which the outer cavity and the inner cavity communicate with each other,
• the plurality of formed from the inner cavity, and disposed along a pressure side blade surface of the blade body and adjacent to each other in tandem in an axially downstream direction,
• the impingement plate having the through holes whose opening density decreases toward the axially downstream side,
• the turbine stationary blade including a cooling hole row which has a plurality of cooling holes communicating with the inner cavity via inlet openings formed in an inner surface of the bottom plate and communicating with a combustion gas flow passage via outlet openings formed in a gas path surface of the bottom plate, the cooling method for the turbine stationary blade, comprising:
• a step of externally supplying cooling air to the outer cavity;
• a step of reducing a pressure of the cooling air, where the cooling air is supplied from the outer cavity to the inner cavity via the through holes formed in the impingement plate disposed in the plurality of cavities, an internal pressure of the cavities decreases toward the axially downstream side, and a differential pressure between the inner cavity and the combustion gas flow passage is maintained constant;
• a step of impingement-cooling an inner surface of the bottom plate with the cooling air;
• and
• a step of film-cooling the gas path surface when the cooling air after the impingement cooling is discharged from the cooling holes formed in the bottom plate to the combustion gas flow passage.