Abstract: A blade is formed with blade air passages connecting the inside of a blade body of the blade and the inside of a platform of the blade. The platform is formed with pressure side passages that extend from the blade air passages toward a circumferential pressure side, are open at a pressure side end face, and are arranged in an axial direction. Further, the platform is formed with a suction side main passage into which cooling air flows, and a suction side passage that communicates with the suction side main passage and extends from the suction side main passage along a suction side end face.

Abstract: A hot part (450) of a gas turbine includes a planar member having a gas path surface (451) that faces at least one of an inner side and an outer side in a radial direction of the rotor in a combustion gas flow path. A cooling flow path is formed in an interior of the planar member along an outer peripheral surface (452) that intersects the gas path surface (451). The planar member has an outer peripheral surface side turbulator (482) that protrudes from an inner surface on an outer peripheral surface (452) side of the cooling flow path.

Abstract: A hot part (450) of a gas turbine includes a planar member having a gas path surface (451) that faces at least one of an inner side and an outer side in a radial direction of the rotor in a combustion gas flow path. A cooling flow path is formed in an interior of the planar member along an outer peripheral surface (452) that intersects the gas path surface (451). The planar member has an outer peripheral surface side turbulator (482) that protrudes from an inner surface on an outer peripheral surface (452) side of the cooling flow path.

Abstract: The thermal load testing device of the present invention makes it possible to produce a large temperature differential between a surface and an interior of a test piece while applying a load to the test piece. A thermal load testing device (1) includes a load applying portion (2) that applies a load to a tubular test piece (10) in an axial line (O) direction, the tubular test piece (10) having a hollow portion (11) that extends along the axial line (O); a cooling fluid supplying portion (3) that causes a cooling fluid to flow through the hollow portion (11); and an infrared image furnace (4) that heats the test piece (10) by a plurality of infrared sources (42) disposed so as to surround the test piece (10) from a whole region in a circumferential direction.

Abstract: A retainer that protrudes from an inner shroud of a vane to a radially inner side and extends in a circumferential direction is formed with an opening that passes through the retainer in an axial direction and defines a space through which air flows. A width of the opening in the circumferential direction is wider than a width of a vane body in the circumferential direction at a radially inner end of the vane body at a position of the retainer in the axial direction.

Abstract: A blade is formed with blade air passages connecting the inside of a blade body of the blade and the inside of a platform of the blade. The platform is formed with a plurality of pressure side passages that extend from the blade air passages toward a circumferential pressure side, are open at a pressure side end face, and are arranged in an axial direction. Further, the platform is formed with a suction side main passage into which cooling air flows, and a suction side passage that communicates with the suction side main passage and extends from the suction side main passage along a suction side end face.

Abstract: An erosion test apparatus includes a combustor configured to obtain a combustion gas by mixing and combusting compressed air and a fuel, and an erodent supply unit configured to supply an erodent to the combustion gas. The erosion test apparatus further includes an accommodation support unit configured to accommodate and support a test piece having a front surface coated through thermal barrier coating, and an accelerator configured to accelerate the combustion gas including the erodent to collide with the test piece.

Abstract: The airfoil member of the present invention is provided with an airfoil body, an end wall which is installed at an end part of the airfoil body, a fillet portion which smoothly connects the end part of the airfoil body with the end wall, and a cooling channel which allows a cooling medium to circulate inside the airfoil body and the end wall and in which two main channels are connected so as to bend in a folding manner at a return channel. The return channel is formed so as to run along the fillet portion on a cross section intersecting with a center line of a profile of the airfoil body and also formed in such a manner that the width thereof in the profile thickness direction is greater than the width of the main channel in the profile thickness direction.

Abstract: The amount of cooling air (cooling medium) can be reduced, and low-temperature cooling air is prevented from being blown out through film cooling holes. Part of a cooling medium impingement-cooling an inner circumferential surface of a blade main body located on a ventral side further impingement-cools the inner circumferential surface of the blade main body located on a dorsal side and is blown out through film cooling holes in the blade main body that are located on the dorsal side.

Abstract: The airfoil member of the present invention is provided with an airfoil body, an end wall which is installed at an end part of the airfoil body in a blade span direction and extends so as to intersect in the blade span direction, a fillet portion which smoothly connects the end part of the airfoil body with the end wall, and a cooling channel which allows a cooling medium to circulate inside the airfoil body and the end wall and in which two main channels extending along the blade span direction are connected so as to bend in a folding manner at a return channel formed on the end wall side. The return channel is formed so as to run along the fillet portion on a cross section intersecting with a center line of a profile of the airfoil body and also formed in such a manner that the width thereof in the profile thickness direction is greater than the width of the main channel in the profile thickness direction.

Abstract: A platform cooling structure for a gas turbine moving blade is provided which is capable of improving cooling performance of a platform and of improving reliability of a moving blade in such a manner that a portion in the vicinity of a side edge of the platform which is away from moving blade cooling passageways and is easily influenced by thermal stress caused by high-temperature combustion gas, that is, an upper surface of the side edge is effectively cooled by guiding high-pressure cooling air, flowing to the moving blade cooling passageways, to a discharge opening formed in a surface of the platform in the vicinity of the side edge of the platform without particularly attaching an additional member such as a cover plate to the platform. A moving blade cooling passageway 17c is formed in the inside of the gas turbine moving blade.

Abstract: The amount of cooling air (cooling medium) can be reduced, and low-temperature cooling air is prevented from being blown out through film cooling holes. Part of a cooling medium impingement-cooling an inner circumferential surface of a blade main body located on a ventral side further impingement-cools the inner circumferential surface of the blade main body located on a dorsal side and is blown out through film cooling holes in the blade main body that are located on the dorsal side.

Abstract: A platform cooling structure for a gas turbine moving blade is provided which is capable of improving cooling performance of a platform and of improving reliability of a moving blade in such a manner that a portion in the vicinity of a side edge of the platform which is away from moving blade cooling passageways and is easily influenced by thermal stress caused by high-temperature combustion gas, that is, an upper surface of the side edge is effectively cooled by guiding high-pressure cooling air, flowing to the moving blade cooling passageways, to a discharge opening formed in a surface of the platform in the vicinity of the side edge of the platform without particularly attaching an additional member such as a cover plate to the platform. A moving blade cooling passageway 17c is formed in the inside of the gas turbine moving blade.

Abstract: The gas turbine stationary blade comprises a stationary blade section provided therein with a passage for cooling air, an inner shroud for supporting the stationary blade section on the side of a discharge port of the cooling air, and a plurality of segments each of which includes at least one stationary blade section and at least one inner shroud. A flow passage is pulled out from the discharge port of the cooling air, and the flow passage is introduced to a front edge corner section of the inner shroud and is extended rearward along a side edge of the inner shroud.

Abstract: The gas turbine stationary blade comprises a stationary blade section provided therein with a passage for cooling air, an inner shroud for supporting the stationary blade section on the side of a discharge port of the cooling air, and a plurality of segments each of which includes at least one stationary blade section and at least one inner shroud. A flow passage is pulled out from the discharge port of the cooling air, and the flow passage is introduced to a front edge corner section of the inner shroud and is extended rearward along a side edge of the inner shroud.

Abstract: The gas turbine stationary blade comprises a stationary blade section provided therein with a passage for cooling air, an inner shroud for supporting the stationary blade section on the side of a discharge port of the cooling air, and a plurality of segments each of which includes at least one stationary blade section and at least one inner shroud. A low passage is pulled out from the discharge port of the cooling air, and the flow passage is introduced to a front edge corner section of the inner shroud and is extended rearward along a side edge of the inner shroud.

Abstract: The gas turbine stationary blade comprises a stationary blade section provided therein with a passage for cooling air, an inner shroud for supporting the stationary blade section on the side of a discharge port of the cooling air, and a plurality of segments each of which includes at least one stationary blade section and at least one inner shroud. A low passage is pulled out from the discharge port of the cooling air, and the flow passage is introduced to a front edge corner section of the inner shroud and is extended rearward along a side edge of the inner shroud.

Abstract: The division wall is made up of a plurality of division wall sections forming a passage wall of high temperature gas which are connected in the direction of arrangement of blades to form a wall surface having a roughly circular cross section as a whole, a gas flow restricting structure for preventing high temperature gas from passing through a gap formed at a connecting portion between the division wall sections in the flow direction of the high temperature gas from the opening on the upstream side of the high temperature gas in the gap, or a gas flow restricting structure for preventing the high temperature gas from being embraced in the gap, for example, a sealing member formed into a prism having a T-shape cross section as a whole composed of a plane portion as a sealing portion and a projected portion for filling the gap is provided.

Abstract: A stationary blade of a gas turbine, which can reduce thermal stress produced at a portion in the vicinity of a rear edge of an inner shroud of the stationary blade. The stationary blade is positioned adjacent to at least one of moving-blade disks in an axial direction of the gas turbine. A concave portion is provided in the inner shroud in a manner such that the concave portion is formed in the vicinity of a rear edge of the inner shroud and on an inner-peripheral face of the inner shroud, where cooling air passes along the inner-peripheral face which faces a rotation shaft of the moving-blade disks; and a protruding portion which protrudes towards the rotation shaft is formed at the rear edge of the inner shroud.

Abstract: The gas turbine stationary blade comprises a stationary blade section provided therein with a passage for cooling air, an inner shroud for supporting the stationary blade section on the side of a discharge port of the cooling air, and a plurality of segments each of which includes at least one stationary blade section and at least one inner shroud. A flow passage is pulled out from the discharge port of the cooling air, and the flow passage is introduced to a front edge corner section of the inner shroud and is extended rearward along a side edge of the inner shroud.