Abstract: In a gas turbine, annular overhang portions are formed on adjacent surfaces of a plurality of rotor discs so as to face each other, surrounding the rotor axis, groove portions are provided circumferentially to the surfaces of the overhang portions facing each other, and sealing structures are annularly installed to the inside of the groove portions. The gas turbine comprises sealing plate assemblies which include the overhang portions, the groove portions and a plurality of plates being formed annularly by being mutually piled up; and detachable retaining members which are provided so as to fix the overhang portions and the sealing plate assemblies together by way of disc engagement portions provided to the overhang portions and sealing plate engagement portions provided to the sealing plate assemblies.

Abstract: A gas turbine, wherein annular overhang portions are formed on adjacent surfaces of a plurality of rotor discs so as to face each other, surrounding the rotor axis, groove portions are provided circumferentially to the surfaces of the overhang portions facing each other, and sealing structures being annularly installed to the inside of the groove portions are provided, comprises sealing plate assemblies which include the overhang portions, the groove portions and a plurality of plates being formed annularly by being mutually piled up; and detachable retaining members which are provided so as to fix the overhang portions and the sealing plate assemblies together by way of disc engagement portions being provided to the overhang portions and sealing plate engagement portions being provided to the sealing plate assemblies.

Abstract: Cooling air passages are formed in the wall of an exhaust casing connected to a turbine casing of a gas turbine. Low pressure air extracted from the low pressure stage of an air compressor of the gas turbine is supplied to the cooling air passage from the portion near the downstream end of the exhaust casing. Cooling air flows through the cooling air passage toward the upstream end of the exhaust casing and then flows into an annular cavity formed in the turbine casing near the portion corresponding to the last turbine stage. Therefore, the metal temperature of the exhaust casing near the upstream end (near the joint between the exhaust casing and the turbine casing) is lowered by the cooling air and, as cooling air of a relatively high temperature is supplied to the cavity in the turbine casing, the metal temperature of the turbine casing near the downstream end becomes higher than that provided by a conventional cooling system.

Abstract: Cooling air passages are formed in the wall of an exhaust casing connected to a turbine casing of a gas turbine. Low pressure air extracted from the low pressure stage of an air compressor of the gas turbine is supplied to the cooling air passage from the portion near the downstream end of the exhaust casing. Cooling air flows through the cooling air passage toward the upstream end of the exhaust casing and then flows into an annular cavity formed in the turbine casing near the portion corresponding to the last turbine stage. Therefore, the metal temperature of the exhaust casing near the upstream end (near the joint between the exhaust casing and the turbine casing) is lowered by the cooling air and, as cooling air of a relatively high temperature is supplied to the cavity in the turbine casing, the metal temperature of the turbine casing near the downstream end becomes higher than that provided by a conventional cooling system.

Abstract: In a central portion of inner tube 28 of combustor 20, pilot fuel nozzle 22 and pilot cone 33 are arranged and main fuel nozzles 21 and main swirlers 32 therearound. Air intake portion (X-1) is provided with rectifier tube 11 for making air intake uniform. In air intake portion (X-2), air holes of appropriate number of pieces are provided in circumferential wall of the inner tube 28. In main swirler portion (X-3) and pilot cone portion (X-4), bolt joint of the main swirlers 32 is employed and optimized welded structure having less influence of thermal stress of the pilot swirler 33 is employed, respectively. Tail tube cooling portion (X-5) is provided with cooling structure having less influence of thermal stress to cool flange 71 portion of tail tube 24 uniformly. By the improvements in the portions (X-1) to (X-5), obstacles in attaining higher temperature in the combustor 20 is dissolved and combustor performance is enhanced.

Abstract: In a central portion of inner tube 28 of combustor 20, pilot fuel nozzle 22 and pilot cone 33 are arranged and main fuel nozzles 21 and main swirlers 32 therearound. Air intake portion (X-1) is provided with rectifier tube 11 for making air intake uniform. In air intake portion (X-2), air holes of appropriate number of pieces are provided in circumferential wall of the inner tube 28. In main swirler portion (X-3) and pilot cone portion (X-4), bolt joint of the main swirlers 32 is employed and optimized welded structure having less influence of thermal stress of the pilot swirler 33 is employed, respectively. Tail tube cooling portion (X-5) is provided with cooling structure having less influence of thermal stress to cool flange 71 portion of tail tube 24 uniformly. By the improvements in the portions (X-1) to (X-5), obstacles in attaining higher temperature in the combustor 20 is dissolved and combustor performance is enhanced.

Abstract: In a central portion of inner tube 28 of combustor 20, pilot fuel nozzle 22 and pilot cone 33 are arranged and main fuel nozzles 21 and main swirlers 32 therearound. Air intake portion (X-1) is provided with rectifier tube 11 for making air intake uniform. In air intake portion (X-2), air holes of appropriate number of pieces are provided in circumferential wall of the inner tube 28. In main swirler portion (X-3) and pilot cone portion (X-4), bolt joint of the main swirlers 32 is employed and optimized welded structure having less influence of thermal stress of the pilot swirler 33 is employed, respectively. Tail tube cooling portion (X-5) is provided with cooling structure having less influence of thermal stress to cool flange 71 portion of tail tube 24 uniformly. By the improvements in the portions (X-1) to (X-5), obstacles in attaining higher temperature in the combustor 20 is dissolved and combustor performance is enhanced.

Abstract: In a steam cooling type gas turbine, a sealing structure for improving sealing performance between an interior of a rotor and a gas path of a turbine section which performs inter-disk sealing such that leakage of cooling steam and self-induced vibration of a baffle plate are prevented. A groove is formed along a circumferential direction in an end face of at least one of disk lands which protrude in opposition to each other between adjacent rotor disks, and an annular sealing member having an interior space is disposed in a sandwiched fashion, being brought into contact under pressure with an inner wall surface of the groove and an end face of the other disk land, or alternatively, an inner wall surface of a groove formed in the other disk land to thereby realize the inter-disk sealing structure for the gas turbine.

Abstract: A steam-cooling gas turbine in which cooling steam is fed from a center portion of a rotating shaft and recovered through passages disposed at the outer side of the center portion with steam leaking through a seal portion from a feeding steam being effectively recovered. Feed steam (30) serving as coolant is supplied to a cavity (27) from a inner cylinder (10) and hence fed to moving blades (11, 12) through recesses (40) formed in a coupling portion (26) interposed between a final-stage disk (24) and a turbine shaft (1), steam feeding pipes (15) and steam feeding passages (11a, 12a). After cooling of the moving blades, steam (31) is recovered through steam recovering passages (11b, 12b), steam recovering pipes (16), radial steam-recovering passages (17), axial steam-recovering passages (18) and outlet openings (5a). The coupling portion (26) is forced to hermetically close under thermal stress, whereby leakage of the feed steam (30) is prevented.

Abstract: A high differential pressure type end rotor seal that is applicable to a steam cooled type gas turbine with the aim to reduce steam leakage to a minimum without being affected by deviation of fins due to thermal elongation. An end rotor rotational seal of a double strip seal type is structured such that fins (3,5) on a stator side (1) and fins (4,6) on a rotor side (2) are arranged in a confronting arrangement and an alternating arrangement. A fin-to-fin pitch (P) is set in a range of 2 to 6 mm and a clearance (C) between apexes of the fins is set in a range of 0.3 to 1.0 mm, thereby a high differential pressure type rotational seal is obtained having minimum steam leakage without being affected by deviation of fins due to thermal elongation between the stator side and the rotor side.

Abstract: A steam-cooled type gas turbine has flange contact surfaces between adjacent rotor disks in which buffering sheets such as gaskets and the like can not be inserted. Further, the contact surfaces have a large diameter, making it difficult or impossible to ensure sufficient surface contact pressure for sealing. Yet preventing the leakage of coolant steam from these flange contact surfaces presents a very important problem. The present invention provides a seal structure for flange contact surfaces of a gas turbine which can enhance the sealability at these portions. The seal structure for flange contact surfaces includes a plurality of labyrinth-like grooves (8) which are formed in at least one of the flange contact surfaces (7). The seal structure of the present invention provides a gas turbine with high reliability and enhanced sealability. Further, a high surface contact pressure is obtained by reducing the area of the flange contact surfaces.

Abstract: A gas turbine having cooling structure disposed on an outside peripheral side in a radial direction of a movable vane in which the cooling efficiency is further improved and sealing structure disposed on an inner peripheral side to suppress leakage of sealing air and prevent invasion of combustion gas into the turbine disc and enlargement of sealing air clearance is prevented. Outwardly in the radial direction, a gap between ring segment (90) and impingement cooling plate (92) disposed on the outer periphery thereof, is divided in the axial direction by a pressure partition plate (97) extending in a circumferential direction so as to provide plural cavities adjusted to different pressures, thereby forming a cooling structure in the ring segment (90). Inside in the radial direction, a static vane inside diameter side cavity (53) and downstream cavity (31) are formed by a box (57) for sealing the static vane inside diameter side cavity (53).

Abstract: A seal device between a fastening bolt and a bolthole in a gas turbine disc reduces cooling steam leaking from a gap between the fastening bolt and the bolthole. A gas turbine disc 31 carrying moving blades, a fastening bolt 33 for fixing the disc 31 passes through a bolthole 32. On a low pressure side of the disc 31, a bolthole diameter enlarged portion 39 having a larger diameter than that of the bolthole 32 is provided. Fastening bolt 33 is provided with a shaft diameter enlarged portion 33a. Seal piece 1 is fitted at one end around the shaft diameter enlarged portion 33a and is engaged at the other end with a stepped portion of the bolthole diameter enlarged portion 39. In operation of the gas turbine, the bolt 33 is biased outwardly be centrifugal force, and the gap between the bolt 33 and the bolthole 32 deforms. Nevertheless, the seal piece 1, having flexibility, deforms at bent portions R.sub.1 and R.sub.

Abstract: A gas turbine moving blade having a cooling medium flow path constructed so that a concave spherical surface (C) is formed on a inside surface of an end portion of the cooling medium path (B) formed in a blade root portion. A hollow pipe (6) is provided in the cooling medium flow path between the blade root portion and a disc. The hollow pipe has a convex spherical surface (D) at one end thereof which engages the concave spherical surface (C) and the other end of the pipe has a convex spherical surface (F) which engages an inside surface of a cooling medium path (E) formed in the disc. The hollow pipe (6) provides communication between the cooling medium path (B) on the blade side and the cooling medium path (E) on the disc side. A supporting structure (7, 8, 9) is provided for holding the hollow pipe in the communicating position.

Abstract: A gas turbine moving blade steam cooling system in which leakage of steam for cooling a moving blade is prevented and thermal stress at a blade root end portion is mitigated. Each end portion of a blade root portion (3) of the moving blade (1) is projected so as to form a projection portions (4a, 4b). A steam passage 5 is provided between the projection portions (4a, 4b) and a steam supply port 5a and a steam recovery port (5b) are provided downwardly with the respect to the steam passage 5. The steam supply port (5a) connects to a steam supply passage (20) and the steam recovery port (5b) connects to a steam recovery passage (21). Steam is supplied from the steam supply port (5a) into a blade interior and is recovered through the steam recovery port (5b). Side surface seal plates (6, 7 and 8) are provided for preventing steam leakage.