Abstract: A collision plate having plural small holes is provided at an interval from a bottom surface of an inner shroud to form a chamber and guides cooling air from the small holes into the chamber. A leading edge flow path is provided at a leading edge side along a width direction and introduces the cooling air. A side flow path is provided along both sides of the inner shroud and guides the cooling air to a trailing edge side. A header is formed along the width direction near the trailing edge and guides the cooling air from the side flow path. Plural trailing edge flow paths are formed at the trailing edge side at intervals along a width direction, in which one end of each flow path is connected to the header and the other end is open at the trailing edge, and the cooling air in the header is ejected from the trailing edge.

Abstract: In a tail tube seal structure of gas turbine, a U-shaped groove is provided at one side of a tail tube seal where a flange of a tail tube outlet is fitted, and a pi-shaped groove is provided at other side of the tail tube seal where a gas pass side flange end is fitted, thereby composing the seal of the connection area. Inclined cooling holes are drilled in the tail tube seal in addition to the cooling holes existing conventionally. The cooling air flows in from the inclined holes and cools the gas pass side of the groove due to the film effect. Therefore, the difference in thermal expansion between the groove and flange end is decreased, the wear of this area is decreased, and the reliability of the seal is enhanced.

Abstract: A turbine blade applicable to a gas turbine has a turbine blade body having film cooling holes, the interior space of which is partitioned into two cavities by a rib. Hollow inserts each having impingement holes are respectively arranged in the cavities to form cooling spaces therebetween. Communication is ensured between the cavities by bypass hole and slits, so that the impingement cooling is interrupted with respect to the prescribed side having a good heat transmission in the turbine blade body. A partition wall is further arranged between the rib and the insert arranged in the trailing-edge side, thus providing a separation between the cooling spaces respectively arranged in the rear side and front side. Thus, it is possible to noticeably reduce the amount of cooling air in the turbine blade body; and it is possible to reduce temperature differences entirely over the turbine blade body as small as possible.

Abstract: It provides a method for managing the remaining life of a high temperature component used in a gas turbine that enables a state of fatigue of a high temperature component to be evaluated and controlled with a high level of accuracy, and a program medium for performing the method for managing the remaining life of a high temperature gas turbine component on a computer. In the method for managing the lifespan of a high temperature component used in a gas turbine, the lifespan of a high temperature component being evaluated is set within an operating area bounded by a component life limit line that is determined on the basis of field data giving the correlation between the operating time and operating cycles for the state of fatigue from past high temperature components. As the program medium, a medium is employed that is provided with a program in which this method is used for managing a lifespan of a high temperature component used in a gas turbine.

Abstract: A base plate and honeycomb member disposed at the inner circumference side of an inside shroud are fixed to the inside shroud, at a phase deviation in the peripheral direction, with respect to the inside shroud, so as to plug the missing range of seal member, out of gaps between adjacent inside shrouds, and therefore leakage of purge air from the missing range of seal member is prevented without adding new constituent members.

Abstract: A collision plate having plural small holes is provided at an interval from a bottom surface of an inner shroud to form a chamber and guides cooling air from the small holes into the chamber. A leading edge flow path is provided at a leading edge side along a width direction and introduces the cooling air. A side flow path is provided along both sides of the inner shroud and guides the cooling air to a trailing edge side. A header is formed along the width direction near the trailing edge and guides the cooling air from the side flow path. Plural trailing edge flow paths are formed at the trailing edge side at intervals along a width direction, in which one end of each flow path is connected to the header and the other end is open at the trailing edge, and the cooling air in the header is ejected from the trailing edge.

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: A turbine blade applicable to a gas turbine has a turbine blade body having film cooling holes, the interior space of which is partitioned into two cavities by a rib. Hollow inserts each having impingement holes are respectively arranged in the cavities to form cooling spaces therebetween. Communication is ensured between the cavities by a communication means, so that the impingement cooling is interrupted with respect to the prescribed side having a good heat transmission in the turbine blade body. A partition wall is further arranged between the rib and the insert arranged in the trailing-edge side, thus providing a separation between the cooling spaces respectively arranged in the rear side and front side. Thus, it is possible to noticeably reduce the amount of cooling air in the turbine blade body; and it is possible to reduce temperature differences entirely over the turbine blade body as small as possible.

Abstract: In a cooling structure for a gas turbine, cooling air diffusion holes are formed from inner surface to outer surface of a platform so as to open from high pressure side blade surface of a moving blade offset in a direction toward low pressure side blade surface of adjacent moving blade confronting the high pressure side blade surface, in a direction of primary flow.

Abstract: In the gas turbine split ring, on an outer peripheral surface 1b between two cabin attachment flanges, a circumferential rib which extends in the circumferential direction and an axial rib which extends in the axial direction and has a height taller than that of the circumferential rib are, respectively, formed in plural lines, so that it is possible to suppress heat deformation in the axial direction which largely contributes to reduction of the tip clearance compared to head deformation in the circumferential direction more efficiently.

Abstract: A gas turbine cooled stationary blade has a blade structure and outer and inner shrouds enhancing cooling efficiency and preventing the occurrence of cracks due to thermal stresses. A blade (1) wall thickness, between 75% and 100% of the blade height of a blade leading edge portion, is made thicker, and the blade (1) wall thickness of other portions is made thinner, as compared with a conventional case. Protruding ribs (4) are provided on a blade (1) convex side inner wall between 0% and 100% of the blade height. A blade (1) trailing edge opening portion is made thinner than the conventional case. Outer shroud (2) is provided with cooling passages (5a, 5b) for air flow in both side end portions. Inner shroud (3) is provided with cooling passages (9a, 9b) for air flow and cooling holes (13a, 13b) for air blow in the side end portions.

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: A base plate and honeycomb member disposed at the inner circumference side of an inside shroud are fixed to the inside shroud, at a phase deviation in the peripheral direction, with respect to the inside shroud, so as to plug the missing range of seal member, out of gaps between adjacent inside shrouds, and therefore leak of purge air from the missing range of seal member is prevented without adding new constituent members.

Abstract: In the gas turbine split ring, on an outer peripheral surface 1b between two cabin attachment flanges, a circumferential rib which extends in the circumferential direction and an axial rib which extends in the axial direction and has a height taller than that of the circumferential rib are, respectively, formed in plural lines, so that it is possible to suppress heat deformation in the axial direction which largely contributes to reduction of the tip clearance compared to head deformation in the circumferential direction more efficiently.

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: In a cooling structure for a gas turbine, cooling air diffusion holes are formed from inner surface to outer surface of a platform so as to open from high pressure side blade surface of a moving blade offset in a direction toward low pressure side blade surface of adjacent moving blade confronting the high pressure side blade surface, in a direction of primary flow.

Abstract: A gas turbine stationary blade having passages (23, 24) that are provided in the stationary blade (10). A front cylindrical insert (2) is provided in the passage (23) and a rear cylindrical insert (5) is provided in the passage (24), and the inserts are supported at two supporting portions (3a, 3b), (6a, 6b), respectively. A projection (1) is provided at a leading edge portion of the blade so that the leading edge, where the thermal loads are high, is made smaller in size and the number of rows of cooling holes (11a) in the leading edge portion is reduced. Air blowing holes (4b) are provided on the dorsal side rear portion of the front insert (2) and film cooling holes (12) are provided on the dorsal side of the blade, both have diameters that are larger than other air blowing and cooling holes provided in the insert (2) and the blade (10), so that dust in the cooling air is caused to flow out, thereby preventing clogging of the holes.

Abstract: Gas turbine cooled stationary blade is improved in the structure of a blade and outer and inner shrouds to enhance cooling efficiency and to prevent occurrence of cracks due to thermal stresses. Blade (1) wall thickness between 75% and 100% of blade height of a blade leading edge portion is made thicker and blade (1) wall thickness of other portions is made thinner, as compared with a conventional case. Protruding ribs (4) are provided on a blade (1) convex side inner wall between 0% and 100% of the blade height. Blade (1) trailing edge opening portion is made thinner than the conventional case. Outer shroud (2) is provided with cooling passages (5a, 5b) for air flow in the shroud both side end portions. Inner shroud (3) is provided with cooling passages (9a, 9b) for air flow and cooling holes (13a, 13b) for air blow in the shroud both side end portions.

Abstract: In gas turbine split rings, end faces having bent surfaces are formed in flanges. Adjoining split rings are coupled together with a groove therebetween to form a cylindrical split ring. Notches are formed in the flanges. These notches are sealed by inserting a seal plate into the notches of adjoining split rings. A hole for passing cooling air is drilled obliquely in the flange. Cooling air is allowed to flow out along the direction of rotation (of the turbine). This cooling air cools the outlet of the groove due to the effect of film cooling. Because of such cooling, high temperature gas is prevented from staying in this area, cooling effect is enhanced, and hence burning of the end portions can be prevented.