Abstract: In order to provide a turbine blade fatigue life evaluating method for quantitively evaluating the fatigue life of a turbine blade, the turbine blade is determined to be within its fatigue life if the creep elongation strain in the longitudinal direction of the turbine blade is less than 0.5% of an initial length, and is determined to exceed its fatigue life if the creep elongation strain is 0.5% or more than the initial length. A turbine blade creep elongation strain measuring apparatus 20 comprises a first fixed end 21, a second fixed end 22, and a dial gauge 24. A dimension in the longitudinal direction is stamped on the surface of a turbine blade.

Abstract: A platform of a gas turbine moving blade suppresses the effects of thermal elongation and thus improves cooling performance. A structure is constituted by a peripheral edge of a platform of a gas turbine moving blade, a bottom of the platform, and a shank of the moving blade. A cavity is blocked by disposing a sealing plate so as to seal the recessed section, while a supply route is formed for supplying air from cooling passages through an interior of the shank to the cavity, each of the passages being for air-cooling the interior of the gas turbine moving blade with air blown out from the cavity to a surface of the platform. A method of appropriately installing the sealing plate is also disclosed.

Abstract: In order to provide a turbine blade fatigue life evaluating method for quantitively evaluating the fatigue life of a turbine blade, the turbine blade is determined to be within its fatigue life if the creep elongation strain in the longitudinal direction of the turbine blade is less than 0.5% of an initial length, and is determined to exceed its fatigue life if the creep elongation strain is 0.5% or more than the initial length. A turbine blade creep elongation strain measuring apparatus 20 comprises a first fixed end 21, a second fixed end 22, and a dial gauge 24. A dimension in the longitudinal direction is stamped on the surface of a turbine blade.

Abstract: An object of the present invention is to provide a ring segment of a gas turbine in which the temperature is maintained low, damage due to high temperature oxidation is prevented, and high temperature deformation is prevented.

Abstract: Holes 38 and 39 have upstream opening portions 38b and 39b and downstream opening portions 38a and 39a which have a larger cross-sectional area than upstream opening portions 38b and 39b, and are formed at top portion TP of each moving blade. Holes 38 and 39 have tapered shapes T1 and T2 or step portions, and preferably, downstream opening portions 38a and 39a are eccentrically formed toward the moving direction. When tip squealer 37 is formed, hole 38 is formed so that its opening portion is provided at the side surface of tip squealer 37. Without covering the holes for cooling which are formed at the top portion of the turbine blade due to rubbing or the like, the turbine blade is accurately cooled and stably driven.

Abstract: The object of the present invention is to provide a turbine moving blade that has high heat resistance and can be used for a long period of time by improving the cooling efficiency of the turbine moving blade, and to improve both the thermal efficiency and operating efficiency of a gas turbine through the use of this turbine moving blade. In order to achieve the object, the present invention provide a turbine moving blade arranged in a combustion gas flow path in which a plurality of blow-out openings for blowing out a cooling medium are formed in its outer surface, wherein among a plurality of cooling medium trailing edge blow-out openings arranged from the vicinity of a blade base to the vicinity of a blade tip along a blade trailing edge of turbine moving blade, the opening area of a blade tip trailing edge blow-out opening located in the vicinity of blade tip is set to be larger than the opening area of the other trailing edge blow-out openings.

Abstract: In order to provide a turbine blade creep life evaluating method for quantitively evaluating the life of a turbine blade, the turbine blade is determined to be within its allowable life if the creep elongation strain in the longitudinal direction of the turbine blade is less than 0.5% of an initial length, and is determined to exceed its allowable life if the creep elongation strain is 0.5% or more than the initial length. A turbine blade creep elongation strain measuring apparatus 20 comprises a first fixed end 21, a second fixed end 22, and a dial gauge 24. A dimension in the longitudinal direction is stamped on the surface of a turbine blade.

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: In order to provide a gas turbine tail tube seal which can prevent an outer shroud and an inner shroud in a first row stationary blade from being damaged by heat and from being worn, and to provide a gas turbine having the above-mentioned gas turbine tail tube seal, in the gas turbine tail tube seal, the width of a combustion gas flow channel in the gas turbine tail tube seal is smaller than the width of a combustion gas flow channel made between an outer surface of the inner shroud and an inner surface of the outer shroud, and the combustion gas flow channel is provided with inclined surfaces having a gradually enlarging cross section toward the first row stationary blade in at least a downstream end section.

Abstract: The object of the present invention is to provide a turbine moving blade that has high heat resistance and can be used for a long period of time by improving the cooling efficiency of the turbine moving blade, and to improve both the thermal efficiency and operating efficiency of a gas turbine through the use of this turbine moving blade. In order to achieve the object, the present invention provide a turbine moving blade arranged in a combustion gas flow path in which a plurality of blow-out openings for blowing out a cooling medium are formed in its outer surface, wherein among a plurality of cooling medium trailing edge blow-out openings arranged from the vicinity of a blade base to the vicinity of a blade tip along a blade trailing edge of turbine moving blade, the opening area of a blade tip trailing edge blow-out opening located in the vicinity of blade tip is set to be larger than the opening area of the other trailing edge blow-out openings.

Abstract: In order to provide a gas turbine tail tube seal which can prevent an outer shroud and an inner shroud in a first row stationary blade from being damaged by heat and from being worn, and to provide a gas turbine having the above-mentioned gas turbine tail tube seal, in the gas turbine tail tube seal, the width of a combustion gas flow channel in the gas turbine tail tube seal is smaller than the width of a combustion gas flow channel made between an outer surface of the inner shroud and an inner surface of the outer shroud, and the combustion gas flow channel is provided with inclined surfaces having a gradually enlarging cross section toward the first row stationary blade in at least a downstream end section.

Abstract: In order to provide a turbine blade fatigue life evaluating method for quantitively evaluating the fatigue life of a turbine blade, the turbine blade is determined to be within its fatigue life if the creep elongation strain in the longitudinal direction of the turbine blade is less than 0.5% of an initial length, and is determined to exceed its fatigue life if the creep elongation strain is 0.5% or more than the initial length. A turbine blade creep elongation strain measuring apparatus 20 comprises a first fixed end 21, a second fixed end 22, and a dial gauge 24. A dimension in the longitudinal direction is stamped on the surface of a turbine blade.

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 gas turbine wherein cooling failure attributable to the occurrence of a horseshoe vortex produced in the vicinity of the stationary blades of the turbine is preventable, the gas turbine including moving blades provided on a rotor side which rotate together with the rotor, and stationary blades provided on a stationary side which cover the periphery of the moving blades and form a combustion gas flow path in the interior, and which are arranged alternately with the moving blades in the rotation axis direction of the rotor.

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: Holes 38 and 39 have upstream opening portions 38b and 39b and downstream opening portions 38a and 39a which have a larger cross-sectional area than upstream opening portions 38b and 39b, and are formed at top portion TP of each moving blade. Holes 38 and 39 have tapered shapes T1 and T2 or step portions, and preferably, downstream opening portions 38a and 39a are eccentrically formed toward the moving direction. When tip squealer 37 is formed, hole 38 is formed so that its opening portion is provided at the side surface of tip squealer 37. Without covering the holes for cooling which are formed at the top portion of the turbine blade due to rubbing or the like, the turbine blade is accurately cooled and stably driven.

Abstract: One object of the present invention is to provide a gas turbine where cooling failure attributable to the occurrence of a horseshoe vortex produced in the vicinity of the stationary blades of the turbine, can be prevented.

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: An object of the present invention is to provide a ring segment of a gas turbine in which the temperature is maintained low, damage due to high temperature oxidization is prevented, and high temperature deformation is prevented.

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.