Mitsubishi Heavy Industries Limited

Mitsubishi Hitachi Power Systems, Ltd.

Tohoku Electric Power Company, Inc.

National Institute for Materials Science

Mitsubishi Power, Ltd.

Mitsubishi Heavy Indsutries, Ltd.

Junichiro Masada

Keizo Tsukagoshi

Kazumasa Takata

Tatsuo Ishiguro

Satoshi Mizukami

Yoshiaki Nishimura

Satoshi Hada

Yuya Fukunaga

Yasuoki Tomita

Masamitsu Kuwabara

Masaki Satoh

Hiroki Takahashi

Norio Oogai

Souji Hasegawa

Masato Hayashi

Shinya Ishikawa

Ikuo Okada

Masaki Taneike

Friedrich Soechting

Hidemichi Koyabu

Hiroshi Harada

Eisaku Ito

Hiroyuki Yamazaki

Taiji Torigoe

Tetsu Konishi

Shunsuke Torii

Toshiharu Kobayashi

Yutaka Koizumi

Yoshitaka Uemura

Tadaharu Yokokawa

Masato Araki

Kyoko Kawagishi

Masao Sakamoto

Koji Watanabe

Shinya Hashimoto

Kiyoshi Suenaga

Hidekazu Uehara

Hiroyuki Kimura

Tanehiro Shinohara

Masanori Yuri

Hitoshi Morimoto

Keita Takamura

Masahiko Mega

Shuji Tanigawa

Yoshifumi Okajima

Hidetaka Oguma

Masaaki Matsuura

Yasuhiko Tsuru

Tomoyuki Hirata

Shigenari Horie

Yoshiaki Yamaguchi

Sachio Shimohata

Daisuke Kudo

Kenichiro Takeishi

Naotoshi Okaya

Tatsuji Takahashi

Yuji Komagome

Ryota Okimoto

Michinari Yuyama

Takuya Suzuki

Yoshiaki Igarashi

Hidataka Oguma

Provided is a high-strength, heat-resistant, Ni-base alloy comprising Co: from 5 to 12%, Cr: from 5 to 12%, Mo: from 0.5 to 3.0%, W: from 3.0 to 6.0%, Al: from 5.5 to 7.2%, Ti: from 1.0 to 3.0%, Ta: from 1.5 to 6.0%, Re: from 0 to 2.0%, and C: from 0.01 to 0.20%. The high-strength, heat-resistant, Ni-base alloy is constituted of a Ni-based alloy, the balance of the Ni-based alloy comprising Ni and inevitable impurities. The density of the high-strength, heat-resistant Ni-base alloy is less than 8.5 g/cm.

High-strength, heat-resistant Ni-base alloy, method for producing same, and gas turbine blade

A rotor shaft has: an inlet flow passage through which gas inside a gas compression flow passage flows into an outer cavity of a downstream-side cavity group; a radial flow passage that provides communication between the outer cavity and an axial communication cavity of the downstream-side cavity group; an axial flow passage that provides communication between the axial communication cavity of the downstream-side cavity group and the axial communication cavity of an upstream-side cavity group; another radial flow passage that provides communication between the axial communication cavity and the outer cavity of the upstream-side cavity group; and an outlet flow passage through which the gas inside the outer cavity of the upstream-side cavity group flows out into the gas compression flow passage.

Compressor rotor, compressor, and gas turbine

A gas turbine combustor comprising a plurality of cooling air passages (A) through which bled pressurized air flows from downstream of a combustion gas flow toward upstream, the plurality of cooling air passages (A) being disposed in a wall portion () side by side and aligned in a flow direction of a combustion gas; wherein the plurality of cooling air passages (A) are divided via a passage transition groove portion () into upstream cooling air passages (A) disposed closer to bled pressurized air inlet holes () and downstream cooling air passages (A) disposed closer to bled pressurized air outlet holes (), and center lines (C) of the upstream cooling air passages are non-collinear with center lines (C) of the downstream cooling air passages.

Combustor cooling structure

Provided is a gas turbine capable of achieving high-speed startup of the gas turbine through quick operation control of an ACC system during startup of the gas turbine, improving the cooling efficiency of turbine stationary components, and quickly carrying out an operation required for cat back prevention during shutdown of the gas turbine. Included are a pressurizing device () connected to a branching channel () branching from the discharge side of a compressor () and capable of carrying out an operation for introducing and pressurizing air independently from the compressor (); a temperature-control-medium supply channel () that guides compressed air pressurized at the pressuring device () to a turbine-cooling-medium channel () provided in stationary components of a turbine (); and a temperature-control-medium return channel () that guides the compressed air that has passed through the turbine-cooling-medium channel () to the discharge side of the compressor () such that the flows are combined, and the pressurizing device () is operated at startup of the gas turbine and in preparation immediately before startup to carry out temperature-raising and cooling by letting the compressed air flow in the turbine-cooling-medium channel (). The pressurizing device () is operated during shutdown of the gas turbine to exhaust the high-temperature gas remaining in the turbine ().

Gas turbine and operating method thereof

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.

Erosion test apparatus, accelerator and erosion test method

Provided is a Ni-based single crystal superalloy containing

Ni-based single crystal superalloy

An inner wall surface of a low-pressure-side side plate opposite to a lateral surface of a thin plate is formed along a direction crossing an axial direction of a rotating shaft such that a gap between the inner wall surface and the thin plate gradually decreases from a radially inner side toward a radially outer side of the low-pressure-side side plate.

Shaft seal mechanism

In a cooling structure for a recovery-type air-cooled gas turbine combustor having a recovery-type air-cooling structure that bleeds, upstream of the combustor, and pressurizes compressed air supplied from a compressor, that uses the bled and pressurized air to cool a wall, and that recovers and reuses the bled and pressurized air as combustion air for burning fuel in the combustor together with a main flow of the compressed air, wall cooling in which cooling air is supplied to cooling air passages formed in the wall of the combustor to perform cooling involves a downstream wall region, closer to a turbine, that is cooled using the bled and pressurized air as the cooling air and an upstream wall region, closer to a burner, that is cooled using, as the cooling air, bled compressed air bled from a main flow of the compressed air through a housing inner space.

Cooling structure for recovery-type air-cooled gas turbine combustor

The Ni-based single crystal alloy disclosed here is a single crystal and has a chemical composition containing, as % by mass,

In a cooling system of ring segment that cools a ring segment of a gas turbine, the segment body of the ring segment is constituted from a collision plate that has a small hole that blows out cooling air, a cooling space that is enclosed by the collision plate and the main body of the segment body; a first cavity that receives the cooling air from the cooling space; and a first cooling passage, of which one end communicates with the first cavity, and the other end blows out the cooling air from openings that are arranged in the side end portion into combustion gas; the openings of the first cooling passages being arranged so that the arrangement pitch of the openings becomes smaller or the opening area of the openings becomes larger on the upstream in the flow direction of the combustion gas than the openings on the downstream, and are arranged so that the arrangement pitch of the openings becomes larger or the opening area of the openings becomes smaller on the downstream in the flow direction of the combustion gas than the openings on the upstream.

Cooling system of ring segment and gas turbine

A cooling system of a ring segment for a gas turbine, which includes first cooling passages disposed in an axial direction of the rotating shaft of a main body of the segment body, second cooling passages disposed at one end portion in a direction approximately perpendicular to the first cooling passages, and blowing a cooling air toward the end portion of a neighboring segment body, and third cooling passages connecting a first cavity, which is disposed approximately perpendicular to the axial direction of the rotating shaft at the upstream-end portion, with a cooling space, which is surrounded by the segment body and a collision plate having a plurality of small holes. The first cooling passages include cooling passages of a first region and a second region having a smaller passage cross-sectional area than the first-region cooling passages or a greater arrangement pitch than the first-region cooling passages.

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.

Stationary blade in gas turbine and gas turbine comprising the same

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Tokyo, Japan

Junichiro Masada