Mitsubishi Hitachi Power Systems, Ltd.

Mitsubishi Heavy Industries Limited

Mitsubishi Power, Ltd.

Hironori Tsukidate

Atsushi Sano

Kunihiro Oga

Hidetoshi Kuroki

Koichi Nishida

Yoshitaka Terada

Kazuo Takahashi

Yasuyuki Watanabe

Takanori Shibata

Naoyuki Nagafuchi

Hidetaro Murata

Nobuaki Kizuka

Yasuo Takahashi

Yasuhiro Horiuchi

Yasunori Kimura

Tatsuya Fukuda

Toshimitsu Morimoto

Hiroyuki Tate

Takeshi Aso

A gas turbine including a turbine driven by a combustion gas, a gas turbine casing that includes an exhaust casing having an inner tube and an outer tube, a bearing that rotatably supports a shaft of the turbine, a bearing casing that holds the bearing, a support leg that supports the gas turbine casing, struts that connect the inner tube and the outer tube, and a first support and a second support that support the bearing casing on the inner tube. The first support is located on a side same as the support leg relative to the struts in a flow direction of the combustion gas. The struts are located between the first support and the second support. The first support is fixed to the inner tube and the bearing casing. The second support is fixed to the inner tube and is in slidable contact with the bearing casing.

Gas turbine with reduced axial displacement under thermal expansion

In a turbine designing and manufacturing method attendant on a material change of a rotor disk of a turbine rotor, a temperature rise time ratio is determined which is a desired ratio of a temperature rise time of the temperature of the rotor disk from a first temperature to a second temperature after the material change to the temperature rise time before the material change. An inter-surface distance between surfaces on upstream and downstream sides of the rotor disk after the material change is determined, and a shape of the rotor disk after the material change is determined based on the inter-surface distance. The turbine is designed based on the determined shape of the rotor disk. After the material change in the shape determined in the designing process, the rotor disk and the turbine are manufactured based on the result of the designing process.

Turbine designing and manufacturing method

The present invention provides a gas turbine nozzle capable of reducing stress related to thermal elongation caused by a rise in gas turbine nozzle temperature and thus reducing stress produced when thermal deformation occurs in the gas turbine nozzle. The gas turbine nozzle according to the present invention includes nozzles formed integrally through an inner perimeter end wall and an outer perimeter end wall. The inner perimeter end wall has an upstream connection portion and a downstream connection portion. The upstream connection portion extends radially inward to be connected to an inner perimeter diaphragm. The downstream connection portion is located downstream from the upstream connection portion and extends radially inward to be connected to the inner perimeter diaphragm. The inner perimeter end wall has a thin-walled portion in a rear edge portion of the inner perimeter end wall, the thin-walled portion corresponding to a reduced wall thickness portion of the rear edge portion of the inner perimeter end wall.

Gas turbine nozzle

The present invention reduces the concentration of thermal stress on a burner. A gas turbine combustor receiving compressed air from a compressor, mixing the compressed air with a fuel, burning the mixture to generate a combustion gas, and supplying the combustion gas to a turbine. The combustor includes: an inner cylinder internally forming a combustion chamber; an outer cylinder covering the inner cylinder and forming a cylindrical outer circumferential flow path between the inner and outer cylinders to allow the compressed air to flow; and a burner mounted on an end of the outer cylinder, which is positioned on an opposite side to a turbine side, and facing the combustion chamber. The burner includes a cylindrical base frame including a cavity distributing the fuel, and fuel nozzles circularly arranged as viewed from the combustion chamber and connected to the cavity. When viewed from the combustion chamber, slits extending radially are formed in the base frame such that each separate the circumferentially adjacent fuel nozzles from each other.

Gas turbine combustor and gas turbine

Provided is a two-shaft gas turbine power generation equipment that includes: an induction motor which receives/transmits power from/to a power system; and a speed reducer that makes a rotational speed of a motor rotor of the induction motor lower than a rotational speed of a compressor rotor of a two-shaft gas turbine. The speed reducer includes: a compressor-side shaft; a motor-side shaft; a first compressor-side helical gear and a second compressor-side helical gear which are mounted on the compressor-side shaft; and a first motor-side helical gear and a second motor-side helical gear which are mounted on the motor-side shaft. The first compressor-side helical gear meshes with the first motor-side helical gear, and the second compressor-side helical gear meshes with the second motor-side helical gear.

Two-shaft gas turbine power generation equipment

The present invention reduces thermal stress that is applied to the weld zone of a cavity formed in the structural material of a burner. The burner includes the cavity distributing a fuel to fuel nozzles. The cavity is demarcated by a groove formed in the structural material of the burner so as to create a level difference at an opening, and a cover fitted into the level difference to close the groove. The cover is formed by a web covering the opening in the groove and a flange extending in the depth direction of the groove to be fitted into the level difference, and is joined by welding to the structural material in such a manner as to have an L-shaped cross-section. When viewed in a cross-section orthogonal to the groove, a cover-side inner surface that is an inner surface of the flange forming a lateral surface of the cavity is flush with a groove-side inner surface that is an inner surface of the groove forming a lateral surface of the cavity. The boundary between the cover-side inner surface and the groove-side inner surface is away from a first corner of the cavity sandwiched between the cover-side inner surface and the inner surface of the web forming a ceiling surface of the cavity.

An object is to reduce the compressor flow rate in comparison with the reference model while maintaining a compression ratio equivalent to that in the reference model. Annulus areas required of a compressorof a derivative gas turbineare determined based on a compressor flow rate and a compression ratio required of the compressorof the derivative gas turbine. Under the condition that the annulus area of each stage of the compressorbecomes equal to the determined annulus area, an inner radius increment and an outer radius decrement of an initial stageare determined, the inner radius increment and the outer radius decrement of each of intermediate stages-are determined so that the inner radius increment is not more than the inner radius increment of the previous stage and the outer radius decrement is not less than the outer radius decrement of the previous stage, and the inner radius increment and the outer radius decrement of a final stageare determined so that the outer radius decrement is not less than the inner radius increment. The compressoris designed by updating design data of components of the reference compressorthat deviated from the specifications due to the determination of the inner radius increment and the outer radius decrement so that the updated design data fulfill the specifications in each of the stages-

Production methods of gas turbine

A regenerative gas turbine combustor is provided that enable to reduce an amount of leakage of compressed air before preheating to compressed air after the preheating. The regenerative gas turbine combustor is such that a compressed air passage between a combustor inner cylinderand a tail cylinder, and a combustor outer cylinderis blocked by a division wallat a position between a bleeding portand an injection port, and compressed air is preheated in a regeneratorand then the compressed air thus preheated is burnt together with fuel. This combustor includes seal rings-, a holderinstalled on the inner circumferential portion of the division walland having ring grooves-for holding the respective seal rings-, and gapsprovided between the inner circumferential surfaces of the ring grooves-and the outer circumferential surfaces of the seal rings-

Sealing mechanism for a regenerative gas turbine combustor

An axial-flow compressor with variable stator vanes located at a single stage can be modified by adding variable stator vanes to the axial-flow compressor so that the variable stator vanes are located at a plurality of stages. The axial-flow compressor includes stator vane rows located at the plurality of stages and include the variable stator vanes extending in a radial direction of the axial-flow compressor and rotating around rotary shafts of the variable stator vanes so as to adjust angles of the variable stator vanes; a plurality of rings are connected to the stator vane rows and drive and rotate the variable stator vanes of the stator vane rows that correspond to respective rings; a plurality of levers that correspond to the plurality of rings; a rotary shaft that holds the plurality of levers so as to enable the levers to pivot.

Axial-flow compressor and modification method

Disclosed is a gas turbine and a gas turbine power facility which flexibly accommodate changes in layout of devices and in user's needs. This invention includes an air compressor; a turbinecoaxially coupled to the compressor; a turbine casingaccommodating the compressorand the turbine; a compressor output shaftfor coupling rotary devices including a power generator, the compressor output shaftprotruding from the turbine casingtowards the compressor; and a turbine output shaftfor coupling the rotary devices, the turbine output shaftprotruding from the turbine casingtowards the turbine

Gas turbine and gas turbine power facilities

It is an object of the present invention to provide a compressor that has a simple structure, produces a damping effect and improves aerodynamic performance. In a compressor including stator bladescircumferentially mounted to an inner circumferential surface side of a casingforming a annular path; and an inner barrelsupported by the casing and arranged on the radially inside of the stator blade as a partition wall on the radial side of the annular path; the stator blade includes an outer shroudmounted to an inner circumferential surface of the casing at a position facing the inner barrel, and an inner shroudsupporting a blade portion at an inner diameter side, the inner shroud being disposed in an annular groove formed in an outer circumferential surface of the inner barrel facing the inner shroud. The stator blade including the outer shroud, the inner shroudand the blade portionsare formed in a monolithic structure by milling.

Growing community of inventors

Yokohama, Japan

Hironori Tsukidate