Abstract: An airfoil including: an airfoil portion having an airfoil surface; and a communication hole extending at least in the airfoil portion and a first opening open in the airfoil surface, through which the first opening is communicated with a second opening provided in a portion of the airfoil. On a cross-section perpendicular to the spanwise direction through a position of the first opening of the spanwise direction, an angle A1 satisfying a condition (a) exists within an angle range ?10 degrees to 10 degrees with respect to an extension line obtained by extending a camber line of the airfoil portion from a leading edge. The condition (a) is a static pressure at a position of the first opening is equal to a static pressure at a position of the second opening when the airfoil portion receives a fluid flow from a direction of the angle A1 toward the leading edge.

Abstract: A blade includes an airfoil portion including a pressure surface and a suction surface each extending between a leading edge and a trailing edge along a spanwise direction, and at least one communication hole extending m the airfoil portion and having a first and a second opening end opening to the pressure and suction surface, respectively. The first opening end is located on a first cross-section perpendicular to the spanwise direction at a first position in the spanwise direction, the second opening end is located on a second cross-section perpendicular to the spanwise direction at a second position in the spanwise direction, and a dimensionless blade chord length position of the first opening end with respect to the leading edge on the first cross-section is larger than a dimensionless blade chord length position of the second opening end with respect to the leading edge on the second cross-section.

Abstract: A gas turbine combustor and a gas turbine; wherein inclined surfaces are provided on inner surfaces of side walls neighboring in a circumferential direction at downstream end portions of transition pieces of combustors, the inclined surfaces being configured to increase a passage area of the transition pieces, a ratio (S/P) is from 0 to 0.2, where (P) is a pitch dimension of first stage vanes, and (S) is a circumferential dimension from an intermediate point between neighboring transition pieces to an upstream end of a first stage vane closest in the circumferential direction; and a ratio (L/P) is from 0.3 to 0.55, where (P) is the pitch dimension, and (L) is an axial dimension from a downstream end of the transition piece to the upstream end of the first stage vane.

Abstract: A rotor blade (22) provided in an axial flow compressor (1) including a rotating shaft, a casing (3), a diffuser portion (4) provided on a downstream side of the casing to communicate with a flow path (C) of the casing and form an annular shape and configured to define a diffuser flow path (DC) in which a cross-sectional area of the flow path expands toward the downstream side, a plurality of stator vane rows (10), and rotor blade rows (20) performing compression of a gas. A plurality of rotor blades are spaced apart from each other in a circumferential direction, and constitute a final rotor blade row (20A) positioned on a most downstream side among the rotor blade rows, and include a blade portion (25) having a larger deflection angle on a hub side and a chip side than at a central portion in a blade height direction.

Abstract: An airfoil includes an airfoil portion having a first airfoil surface and a second airfoil surface respectively extending along a spanwise direction between a leading edge and a trailing edge and having a symmetrical shape with respect to a chord, and at least one communication hole extending in the airfoil portion and having a first opening end opening to the first airfoil surface and a second opening end opening to the second airfoil surface. The first opening end is located on a first cross-section orthogonal to the spanwise direction. The second opening end is located on a second cross-section orthogonal to the spanwise direction. In the first cross-section or the second cross-section, an angle A1 exists within an angle range of ?10 degrees or more and 10 degrees or less with reference to a straight line centered on the leading edge and parallel to an inflow direction of fluid.

Abstract: A blade includes an airfoil portion including a pressure surface and a suction surface each extending between a leading edge and a trailing edge along a spanwise direction, and at least one communication hole extending m the airfoil portion and having a first and a second opening end opening to the pressure and suction surface, respectively. The first opening end is located on a first cross-section perpendicular to the spanwise direction at a first position in the spanwise direction, the second opening end is located on a second cross-section perpendicular to the spanwise direction at a second position in the spanwise direction, and a dimensionless blade chord length position of the first opening end with respect to the leading edge on the first cross-section is larger than a dimensionless blade chord length position of the second opening end with respect to the leading edge on the second cross-section.

Abstract: An airfoil including: an airfoil portion having an airfoil surface; and a communication hole extending at least in the airfoil portion and a first opening open in the airfoil surface, through which the first opening is communicated with a second opening provided in a portion of the airfoil. On a cross-section perpendicular to the spanwise direction through a position of the first opening of the spanwise direction, an angle A1 satisfying a condition (a) exists within an angle range ?10 degrees to 10 degrees with respect to an extension line obtained by extending a camber line of the airfoil portion from a leading edge. The condition (a) is a static pressure at a position of the first opening is equal to a static pressure at a position of the second opening when the airfoil portion receives a fluid flow from a direction of the angle A1 toward the leading edge.

Abstract: A gas turbine provided with an air bleeder tube (1) that, during startup, bleeds a portion of the compressed air of a compressor from the compressor and discharged the bled air into a cylindrical exhaust duct (20), wherein the air bleeder tube (1) is disposed at a portion that does not obstruct the flow of the main flow of combustion gas.

Abstract: A blade or vane row including a plurality of blade or vane segments (10), each including a blade or vane body (11), being disposed in a rotational direction to form an annular shape and defining a flow path (R) along which working fluid (G) flows. The blade or vane segments (10) each have an end wall portion (12, 13) that faces the flow path (R) and extends in a rotational direction. The blade or vane body extends from the end wall portion, and the blade or vane body is disposed in the flow path (R). Surfaces (15A, 15B) of the end wall portion (12, 13) in the rotational direction each include an inclined portion (16). The inclined portion is inclined so as to extend to a flow path (R) side and downstream in a turn direction of the working fluid (G) that flows along the flow path (R).

Abstract: A gas turbine combustor and a gas turbine; wherein inclined surfaces are provided on inner surfaces of side walls neighboring in a circumferential direction at downstream end portions of transition pieces of combustors, the inclined surfaces being configured to increase a passage area of the transition pieces, a ratio (S/P) is from 0 to 0.2, where (P) is a pitch dimension of first stage vanes, and (S) is a circumferential dimension from an intermediate point between neighboring transition pieces to an upstream end of a first stage vane closest in the circumferential direction; and a ratio (L/P) is from 0.3 to 0.55, where (P) is the pitch dimension, and (L) is an axial dimension from a downstream end of the transition piece to the upstream end of the first stage vane.

Abstract: A rotor blade (22) provided in an axial flow compressor (1) including a rotating shaft, a casing (3), a diffuser portion (4) provided on a downstream side of the casing to communicate with a flow path (C) of the casing and form an annular shape and configured to define a diffuser flow path (DC) in which a cross-sectional area of the flow path expands toward the downstream side, a plurality of stator vane rows (10), and rotor blade rows (20) performing compression of a gas. A plurality of rotor blades are spaced apart from each other in a circumferential direction, and constitute a final rotor blade row (20A) positioned on a most downstream side among the rotor blade rows, and include a blade portion (25) having a larger deflection angle on a hub side and a chip side than at a central portion in a blade height direction.

Abstract: In a turbine exhaust structure and a gas turbine, a turbine casing (26) formed in an annular shape to constitute a combustion gas passage A is provided. An exhaust diffuser (31) formed in an annular shape to constitute a flue gas passage (B) is connected to the turbine casing (26) to constitute a configuration. By providing a pressure loss body (61) in the exhaust diffuser (31), efficient pressure recovery can be carried out, which improves turbine efficiency, thereby enabling improvement of the performance.

Abstract: A flow path structure includes a wall surface in which a flow path is formed, a structure configured to extend in a direction intersecting a main stream direction of a fluid flowing through the flow path from the wall surface, and a concave section forming region formed throughout a range including the structure in the main stream direction and having a concave section formed in the wall surface. As the structure occupies a partial range of the flow path in the flow path cross section intersecting the main stream, a cross-sectional area of the flow path is varied in accordance with positional variation in the main stream direction.

Abstract: A gas turbine burns the air compressed in a compressor with supplying fuel in a combustor so as to obtain rotary power by supplying the generated combustion gas to a turbine. The turbine includes turbine vane elements and turbine blade elements that are alternately positioned in a direction in which the combustion gas fluidizes in a turbine cylinder having a cylindrical shape, and a flue gas diffuser having a cylindrical shape and connected to a rear portion of the turbine cylinder. The turbine blade element includes a plurality of turbine blades positioned at equal intervals in the circumference direction. The turbine blades have a throat width on a longitudinal end side made larger than a throat width on a longitudinally intermediate portion side. This efficiently restores the pressure of the flue gas. This improves the efficiency of the turbine so that the performance can be improved.

Abstract: A gas turbine provided with an air bleeder tube (1) that, during startup, bleeds a portion of the compressed air of a compressor from the compressor and discharged the bled air into a cylindrical exhaust duct (20), wherein the air bleeder tube (1) is disposed at a portion that does not obstruct the flow of the main flow of combustion gas.

Abstract: A gas turbine includes a turbine blade, a turbine vane, a ring segment circumferentially surrounding the turbine blade, an outer shroud circumferentially surrounding the turbine vane, and a combustion gas flow-path provided in the ring segment and the outer shroud. The outer shroud is positioned on an upstream side of the ring segment in a gas flow direction of the combustion gas. Seal gas, of which temperature is lower than that of the combustion gas, is fed between the ring segment- and the outer shroud into the combustion gas flow-path. The outer shroud has a guide surface that is provided on an inner circumference thereof on a downstream side of the gas flow direction. The guide surface is formed such that a flow passage area of the combustion gas flow-path is gradually increased.

Abstract: A gas turbine provided with an air bleeder tube (1) that, during startup, bleeds a portion of the compressed air of a compressor from the compressor and discharged the bled air into a cylindrical exhaust duct (20), wherein the air bleeder tube (1) is disposed at a portion that does not obstruct the flow of the main flow of combustion gas.

Abstract: A blade or vane row including a plurality of blade or vane segments (10), each including a blade or vane body (11), being disposed in a rotational direction to form an annular shape and defining a flow path (R) along which working fluid (G) flows. The blade or vane segments (10) each have an end wall portion (12, 13) that faces the flow path (R) and extends in a rotational direction. The blade or vane body extends from the end wall portion, and the blade or vane body is disposed in the flow path (R). Surfaces (15A, 15B) of the end wall portion (12, 13) in the rotational direction each include an inclined portion (16). The inclined portion is inclined so as to extend to a flow path (R) side and downstream in a turn direction of the working fluid (G) that flows along the flow path (R).

Abstract: In a communicating structure between combustors that generates combustion gas inside pipe pieces and a turbine portion that generates a rotational driving force by making the combustion gas sequentially pass through a turbine stage formed of turbine stator vanes and turbine rotor blades, at least some of the first-stage turbine stator vanes closest to the combustor among the turbine stator vanes are disposed downstream of sidewalls of one pipe piece and another pipe piece that are adjacent to each other, and the distance from leading edges of the first-stage turbine stator vanes disposed downstream of the sidewalls of the pipe pieces to end portions of the sidewalls closer to the turbine portion is equal to or less than a spacing between an internal surface of the sidewall of the one pipe piece and an internal surface of the sidewall of the other pipe piece.

Abstract: A one-stage stator vane cooling structure consisting of: a plurality of linking members that are provided between a plurality of combustors disposed in the circumferential direction of a gas turbine; and cooling holes that are provided in each of the one-stage stator vanes to discharge a cooling gas from the inside to the outside of the one-stage stator vanes in order to cool the one-stage stator vanes on the periphery of the stagnation line of the combustion gas flowing in from the plurality of combustors; in which the cooling holes are formed at positions that are determined according to the relative positions of the one-stage stator vanes and the linking members disposed near the one-stage stator vanes.