Abstract: A platform of a gas turbine rotor blade according to one embodiment includes a groove portion recessed from an end surface on a trailing edge side toward a leading edge side. A bottom portion of the groove portion overlaps at least a blade-shaped portion when viewed from a radial direction.

Abstract: A platform of a gas turbine rotor blade according to one embodiment includes a groove portion recessed from an end surface on a trailing edge side toward a leading edge side. A bottom portion of the groove portion overlaps at least a blade-shaped portion when viewed from a radial direction.

Abstract: A turbine blade includes: a pin fin passage formed inside a trailing edge portion of an airfoil part, extending toward a trailing edge of the airfoil part, and opening to outside of the airfoil part at the trailing edge; and a plurality of pin fins connecting a pair of facing inner walls constituting the pin fin passage. The pin fin passage includes a first region and a second region on the trailing edge side of the first region. The plurality of pin fins includes a plurality of first pin fins disposed in the first region and a plurality of second pin fins disposed in the second region. A first diameter of the first pin fins is larger than a second diameter of the second pin fins. A first pin pitch of the first pin fins is larger than a second pin pitch of the second pin fins.

Abstract: A turbine stator vane includes: a blade body having an outer end and an inner end in a blade height direction; a cooling passage extending along the blade height direction inside the blade body; an outer shroud connected to the blade body at the outer end side of the blade body in the blade height direction; an inner shroud connected to the blade body at the inner end side of the blade body in the blade height direction; and an internal passage forming part extending along the blade height direction inside the cooling passage. The internal passage forming part forms an internal passage communicating with each of an outer space on an opposite side to the blade body across the outer shroud and an inner space on an opposite side to the blade body across the inner shroud.

Abstract: A turbine blade includes: a blade body; and a plurality of cooling passages each extending inside the blade body along a blade height direction and are connected to each other via return portions disposed at an end portion in the blade height direction to form a serpentine passage. The cooling passages include: an upstream passage disposed on an upstream side in a flow of a cooling fluid among the plurality of cooling passages; a downstream passage disposed on a downstream side in the flow of the cooling fluid among the plurality of cooling passages; and at least one intermediate passage disposed between the upstream passage and the downstream passage among the plurality of cooling passages. The upstream passage is disposed closest to a leading edge in a chord direction of the blade body among the plurality of cooling passages formed inside the blade body and extending along the height direction.

Abstract: According to an embodiment, a turbine casing comprises a cooling air supply unit configured to supply a cooling air to an interior space of a casing of a gas turbine, and the cooling air supply unit includes: a first supply unit disposed in an upper half of the casing so as to face a first region and configured to supply the cooling air to the first region, where the first region is a region on a radially outer side of a plurality of combustors arranged annularly around a rotor; and a second supply unit disposed so as to face a second region and configured to supply the cooling air to the second region, where the second region is a region on a radially inner side of the plurality of combustors.

Abstract: An optical multilayer film of a light-shielding member includes light absorbing layers that absorb visible light and dielectric layers that are made of a dielectric such that a total number of the layers is 4 or more. A surface-side light absorption thickness which is a total of physical film thicknesses of the light absorbing layers disposed between the outermost layer and a next outermost layer is not less than 6 nm and not greater than 17 nm. A base-side light absorption thickness which is a total of physical film thicknesses of the light absorbing layers disposed between the next outermost layer and the base is not less than 60 nm. A specific proportion, regarding a specific surface layer thickness which is a total of the physical film thicknesses of layers from a base-side maximum-thickness light absorbing layer to the outermost layer, is not less than 0.34.

Abstract: A turbine blade includes an airfoil portion having a hollow portion extending along the blade height direction and film cooling holes arranged along the blade height direction, and an insert disposed in the hollow portion along the blade height direction and having impingement cooling holes. The insert includes a first high-density opening region having a higher density of the impingement cooling holes than in other surface regions of the insert. The geometric center of the first high-density opening region is positioned on the suction side of the airfoil portion in the leading-edge-to-trailing-edge direction and on the outer side of the midpoint of the airfoil portion in the blade height direction. The film cooling holes includes upstream film cooling holes disposed in the suction surface of the airfoil portion at a position corresponding to the first high-density opening region or at a position closer to a leading edge than the position.

Abstract: A turbine blade, a gas turbine, an intermediate product of the turbine blade, and a method of manufacturing the turbine blade are disclosed. The turbine blade has a blade body having hollow shape, cavities defined inside the blade body, and a cooling passage that opens from the cavities to a rear end portion of the blade body. The cooling passage includes: a first passage on a third cavity side and having a width that becomes narrower from the third cavity side toward the rear end portion of the blade body; and a second passage on a rear end portion side of the blade body and having a width that is constant from the third cavity side toward the rear end portion of the blade body.

Abstract: A first-stage stator vane for a gas turbine includes: a first portion partially defining an airfoil which includes a pressure surface, a suction surface, and a trailing edge; and a second portion positioned at a leading-edge side of the airfoil with respect to the first portion, the second portion having a recess portion or a protruding portion. The recess portion or the protruding portion has a pair of side wall surfaces, and an angle formed between the pair of side wall surfaces is less than 90 angular degrees.

Abstract: A turbine rotor blade includes: a blade body including a pressure surface and a suction surface; and a tip shroud on a tip portion of the blade body, the tip shroud being inclined outward in a radial direction from the pressure surface to the suction surface in an axial direction. The tip shroud includes a fin at a center portion in a circumferential direction, the fin extending radially outward, a pressure-side tip shroud on the pressure surface side, and a suction-side tip shroud. The suction-side tip shroud includes a suction-side contact block at a front edge end of the tip shroud. The pressure-side tip shroud includes a pressure-side contact block at a rear edge end of the tip shroud, the suction-side contact block includes a first surface facing in the circumferential direction, and the pressure-side contact block includes a second surface facing in a direction opposite to the circumferential direction.

Abstract: According to an embodiment, a turbine casing comprises a cooling air supply unit configured to supply a cooling air to an interior space of a casing of a gas turbine, and the cooling air supply unit includes: a first supply unit disposed in an upper half of the casing so as to face a first region and configured to supply the cooling air to the first region, where the first region is a region on a radially outer side of a plurality of combustors arranged annularly around a rotor; and a second supply unit disposed so as to face a second region and configured to supply the cooling air to the second region, where the second region is a region on a radially inner side of the plurality of combustors.

Abstract: A gas turbine includes a plurality of first-stage stator vanes arranged in a circumferential direction, the first-stage stator vanes each including a vane surface having a pressure surface and a suction surface, a first combustor disposed on a suction surface side of one of the first-stage stator vanes, the first combustor having a first combustor outlet which includes a first side wall portion extending along a radial direction, and a second combustor disposed on a pressure surface side of the one of the first-stage stator vanes and adjacent to the first combustor in the circumferential direction. The one of the first-stage stator vanes satisfies 0.05??y/P?0.25, where ?y is a protruding amount of the suction surface from the inner wall surface of the first side wall portion to the circumferential direction, and P is an arrangement pitch of the first-stage stator vanes in the circumferential direction.

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 first-stage stator vane for a gas turbine includes: a first portion partially forming an airfoil which includes a pressure surface, a suction surface, and a trailing edge; and a second portion positioned at a leading-edge side of the airfoil with respect to the first portion, the second portion having a recess portion or a protruding portion. The recess portion or the protruding portion of the second portion has a pair of side wall surfaces, and an angle formed between the pair of side wall surfaces is less than 90 angular degrees.

Abstract: A turbine rotor blade includes: a blade body including a pressure surface and a suction surface; and a tip shroud on a tip portion of the blade body, the tip shroud being inclined outward in a radial direction from the pressure surface to the suction surface in an axial direction. The tip shroud includes a fin at a center portion in a circumferential direction, the fin extending radially outward, a pressure-side tip shroud on the pressure surface side, and a suction-side tip shroud. The suction-side tip shroud includes a suction-side contact block at a front edge end of the tip shroud. The pressure-side tip shroud includes a pressure-side contact block at a rear edge end of the tip shroud, the suction-side contact block includes a first surface facing in the circumferential direction, and the pressure-side contact block includes a second surface facing in a direction opposite to the circumferential direction.

Abstract: A turbine blade includes an airfoil portion having a hollow portion extending along the blade height direction and film cooling holes arranged along the blade height direction, and an insert disposed in the hollow portion along the blade height direction and having impingement cooling holes. The insert includes a first high-density opening region having a higher density of the impingement cooling holes than in other surface regions of the insert. The geometric center of the first high-density opening region is positioned on the suction side of the airfoil portion in the leading-edge-to-trailing-edge direction and on the outer side of the midpoint of the airfoil portion in the blade height direction. The film cooling holes includes upstream film cooling holes disposed in the suction surface of the airfoil portion at a position corresponding to the first high-density opening region or at a position closer to a leading edge than the position.

Abstract: A gas turbine includes a plurality of first-stage stator vanes arranged in a circumferential direction, the first-stage stator vanes each including a vane surface having a pressure surface and a suction surface, a first combustor disposed on a suction surface side of one of the first-stage stator vanes, the first combustor having a first combustor outlet which includes a first side wall portion extending along a radial direction, and a second combustor disposed on a pressure surface side of the one of the first-stage stator vanes and adjacent to the first combustor in the circumferential direction. The one of the first-stage stator vanes satisfies 0.05??y/P?0.25, where ?y is a protruding amount of the suction surface from the inner wall surface of the first side wall portion to the circumferential direction, and P is an arrangement pitch of the first-stage stator vanes in the circumferential direction.

Abstract: A turbine vane (3) includes: a vane body (21); a plate-like inner shroud (22) provided at a radially inner end of the vane body (21); and a plate-like outer shroud (23) provided at a radially outer end of the vane body (21). The vane body (21) includes a serpentine channel (30) which is formed so as to meander inside the vane body (21) in the radial direction and through which a cooling medium flows. The inner shroud (22) includes a cooling path (40) which has one end open at the downstream end side of the serpentine channel (30) and the other end open at a trailing edge (22D) of the inner shroud (22) and through which the serpentine channel (30) communicates with the outside of the inner shroud (22).

Abstract: Provided are a turbine blade, a gas turbine, an intermediate product of the turbine blade, and a method of manufacturing the turbine blade. This turbine blade has a blade body having hollow shape, cavities provided inside the blade body, and a cooling passage that opens from the cavities to the rear end portion of the blade body. The cooling passage includes: a first passage provided on the third cavity side and having a width that becomes narrower from the third cavity side toward the rear end portion of the blade body; and a second passage provided on the rear end portion side of the blade body and having a width that is constant from the third cavity side toward the rear end portion of the blade body.