Mitsubishi Heavy Industries Limited

Mitsubishi Power, Ltd.

Mitsubishi Hitachi Power Systems, Ltd.

National Institute for Materials Science

Tohoku Electric Power Company, Inc.

Shunsuke Torii

Yasuoki Tomita

Masamitsu Kuwabara

Satoshi Hada

Masanori Yuri

Friedrich Soechting

Kiyoshi Suenaga

Tatsuo Ishiguro

Keisuke Matsuyama

Ryotaro Magoshi

Kenji Sato

Satoshi Mizukami

Junichiro Masada

Kentaro Tokuyama

Hiroshi Harada

Hiroyuki Yamamoto

Shinya Hashimoto

Takashi Fujii

Hideaki Sugishita

Satoshi Tanimura

Hiroyuki Aoki

Eisaku Ito

Shinya Ishikawa

Ikuo Okada

Tetsu Konishi

Sunao Aoki

Masaaki Matsuura

Tetsuya Shimmyo

Hidemichi Koyabu

Tatsuaki Fujikawa

Naoya Tatsumi

Jun Kubota

Yasuo Miyahisa

Shinichi Inoue

Keita Takamura

Koichiro Iida

Toshiharu Kobayashi

Keizo Tsukagoshi

Yasuro Sakamoto

Yukihiro Hashimoto

Yutaka Koizumi

Yoshitaka Uemura

Yoshifumi Okajima

Hiroaki Kishida

Hidetaka Oguma

Saki Matsuo

Tadaharu Yokokawa

Masaki Taneike

Daigo Fujimura

Kyoko Kawagishi

Kazumasa Takata

Shigehiro Shiozaki

Norifumi Hirata

Yuya Fukunaga

Shingo Yamasue

Hitoshi Kitagawa

Yoshiaki Nishimura

Yoshiyuki Morii

Sachio Shimohata

Yusuke Izumi

Charles Ellis

Hisato Arimura

Hiroki Kitada

Yoshimasa Takaoka

Akihiro Murakami

Michinari Yuyama

Kotaro Ohshima

Yamato Shiramasa

Ryota Okimoto

Akihiko Shirota

Masao Sakamoto

Takaya Koda

Mark J Wagner

Yuji Komagome

Ryo Murata

Yoshiaki Igarashi

Yuko Suenaga

A system includes a hydrogen gas production system and a power generation system. The hydrogen gas production system includes a heated gas supply line configured for flow of a heated gas, a hydrocarbon supply line, a catalytic pyrolysis reactor configured to be in thermal contact with the heated gas of the heated gas supply line and produce a hydrogen containing gas by pyrolyzing a hydrocarbon introduced therein via the hydrocarbon supply line, and a separator configured to extract a hydrogen gas from the hydrogen containing gas discharged from the catalytic pyrolysis reactor. The power generation system includes a heated gas collection line configured to collect the heated gas after the thermal contact with the catalytic pyrolysis reactor and supply the heated gas to the power generation system, and a gas turbine having a combustor configured to burn the hydrogen gas introduced therein from the separator via a hydrogen supply line.

System and method for gas turbine integrated hydrogen production with zero carbon dioxide emissions

A turbine blade includes: a pin fin passage formed inside a trailing edge portion of an airfoil part, extending toward a trailing edge of the airfoil part, and opening to outside of the airfoil part at the trailing edge; and a plurality of pin fins connecting a pair of facing inner walls constituting the pin fin passage. The pin fin passage includes a first region and a second region on the trailing edge side of the first region. The plurality of pin fins includes a plurality of first pin fins disposed in the first region and a plurality of second pin fins disposed in the second region. A first diameter of the first pin fins is larger than a second diameter of the second pin fins. A first pin pitch of the first pin fins is larger than a second pin pitch of the second pin fins.

Turbine blade and gas turbine

A turbine blade includes: a blade body; and a plurality of cooling passages each extending inside the blade body along a blade height direction and are connected to each other via return portions disposed at an end portion in the blade height direction to form a serpentine passage. The cooling passages include: an upstream passage disposed on an upstream side in a flow of a cooling fluid among the plurality of cooling passages; a downstream passage disposed on a downstream side in the flow of the cooling fluid among the plurality of cooling passages; and at least one intermediate passage disposed between the upstream passage and the downstream passage among the plurality of cooling passages. The upstream passage is disposed closest to a leading edge in a chord direction of the blade body among the plurality of cooling passages formed inside the blade body and extending along the height direction.

A platform of a gas turbine rotor blade according to one embodiment includes a groove portion recessed from an end surface on a trailing edge side toward a leading edge side. A bottom portion of the groove portion overlaps at least a blade-shaped portion when viewed from a radial direction. When a pressure-side end portion of the platform with respect to the bottom portion is defined as a first point, a tangent line that is a tangent line with respect to the bottom portion, extending along a plane intersecting the radial direction at the first point is defined as a first tangent line, when an intersection between a line segment connecting the bottom portion and an end portion on the trailing edge side of a serpentine cooling flow path provided inside the blade-shaped portion when viewed from the radial direction and an end portion on the trailing edge side of the blade-shaped portion is defined as a second point, a tangent line that is a tangent line with respect to the bottom portion, extending along the plane at the second point is defined as a second tangent line, and when an intersection between the first tangent line and the second tangent line when viewed from the radial direction is defined as a third point. As viewed from the radial direction, the third point exists on a side opposite to the end portion on the trailing edge side of the blade-shaped portion across a straight line connecting the first point and the second point.

Gas turbine rotor blade and gas turbine

According to an embodiment, a turbine casing comprises a cooling air supply unit configured to supply a cooling air to an interior space of a casing of a gas turbine, and the cooling air supply unit includes: a first supply unit disposed in an upper half of the casing so as to face a first region and configured to supply the cooling air to the first region, where the first region is a region on a radially outer side of a plurality of combustors arranged annularly around a rotor; and a second supply unit disposed so as to face a second region and configured to supply the cooling air to the second region, where the second region is a region on a radially inner side of the plurality of combustors.

Turbine casing, gas turbine, and method for preventing deformation of turbine casing

A turbine blade includes an airfoil portion having a hollow portion extending along the blade height direction and film cooling holes arranged along the blade height direction, and an insert disposed in the hollow portion along the blade height direction and having impingement cooling holes. The insert includes a first high-density opening region having a higher density of the impingement cooling holes than in other surface regions of the insert. The geometric center of the first high-density opening region is positioned on the suction side of the airfoil portion in the leading-edge-to-trailing-edge direction and on the outer side of the midpoint of the airfoil portion in the blade height direction. The film cooling holes includes upstream film cooling holes disposed in the suction surface of the airfoil portion at a position corresponding to the first high-density opening region or at a position closer to a leading edge than the position.

A turbine blade, a gas turbine, an intermediate product of the turbine blade, and a method of manufacturing the turbine blade are disclosed. The turbine blade has a blade body having hollow shape, cavities defined inside the blade body, and a cooling passage that opens from the cavities to a rear end portion of the blade body. The cooling passage includes: a first passage on a third cavity side and having a width that becomes narrower from the third cavity side toward the rear end portion of the blade body; and a second passage on a rear end portion side of the blade body and having a width that is constant from the third cavity side toward the rear end portion of the blade body.

Turbine blade, gas turbine, intermediate product of turbine blade, and method of manufacturing turbine blade

A first-stage stator vane for a gas turbine includes: a first portion partially defining an airfoil which includes a pressure surface, a suction surface, and a trailing edge; and a second portion positioned at a leading-edge side of the airfoil with respect to the first portion, the second portion having a recess portion or a protruding portion. The recess portion or the protruding portion has a pair of side wall surfaces, and an angle formed between the pair of side wall surfaces is less than 90 angular degrees.

First-stage stator vane for gas turbine, gas turbine, stator vane unit for gas turbine, and combustor assembly

A turbine rotor blade includes: a blade body including a pressure surface and a suction surface; and a tip shroud on a tip portion of the blade body, the tip shroud being inclined outward in a radial direction from the pressure surface to the suction surface in an axial direction. The tip shroud includes a fin at a center portion in a circumferential direction, the fin extending radially outward, a pressure-side tip shroud on the pressure surface side, and a suction-side tip shroud. The suction-side tip shroud includes a suction-side contact block at a front edge end of the tip shroud. The pressure-side tip shroud includes a pressure-side contact block at a rear edge end of the tip shroud, the suction-side contact block includes a first surface facing in the circumferential direction, and the pressure-side contact block includes a second surface facing in a direction opposite to the circumferential direction.

Turbine rotor blade, turbo machine, and contact surface manufacturing method

A gas turbine includes a plurality of first-stage stator vanes arranged in a circumferential direction, the first-stage stator vanes each including a vane surface having a pressure surface and a suction surface, a first combustor disposed on a suction surface side of one of the first-stage stator vanes, the first combustor having a first combustor outlet which includes a first side wall portion extending along a radial direction, and a second combustor disposed on a pressure surface side of the one of the first-stage stator vanes and adjacent to the first combustor in the circumferential direction. The one of the first-stage stator vanes satisfies 0.05&#x2264;&#x394;y/P&#x2264;0.25, where &#x394;y is a protruding amount of the suction surface from the inner wall surface of the first side wall portion to the circumferential direction, and P is an arrangement pitch of the first-stage stator vanes in the circumferential direction.

Gas turbine including first-stage stator vanes

A gas turbine combustor and a gas turbine; wherein inclined surfaces are provided on inner surfaces of side walls neighboring in a circumferential direction at downstream end portions of transition pieces of combustors, the inclined surfaces being configured to increase a passage area of the transition pieces, a ratio (S/P) is from 0 to 0.2, where (P) is a pitch dimension of first stage vanes, and (S) is a circumferential dimension from an intermediate point between neighboring transition pieces to an upstream end of a first stage vane closest in the circumferential direction; and a ratio (L/P) is from 0.3 to 0.55, where (P) is the pitch dimension, and (L) is an axial dimension from a downstream end of the transition piece to the upstream end of the first stage vane.

Gas turbine combustor transition piece including inclined surface at downstream end portions for reducing pressure fluctuations

A turbine vane () includes: a vane body (); a plate-like inner shroud () provided at a radially inner end of the vane body (); and a plate-like outer shroud () provided at a radially outer end of the vane body (). The vane body () includes a serpentine channel () which is formed so as to meander inside the vane body () in the radial direction and through which a cooling medium flows. The inner shroud () includes a cooling path () which has one end open at the downstream end side of the serpentine channel () and the other end open at a trailing edge (D) of the inner shroud () and through which the serpentine channel () communicates with the outside of the inner shroud ().

Turbine vane, turbine, and turbine vane modification method

Provided is a Ni-based single crystal superalloy containing

Ni-based single crystal superalloy

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

A blade cooling structure of a gas turbine, which can reduce the pressure loss of a cooling medium without decreasing the heat transfer coefficient, is provided. The blade cooling structure comprises a cooling passage () for flowing cooling air (A) from a proximal end portion () toward a blade portion () of a moving blade (), a plurality of turbulators () arranged, on both wall surfaces of the cooling passage () opposing each other, in such a manner as to be inclined with respect to the flowing direction of the cooling air (A), and a plurality of dimples () formed in a downstream region (N) downstream of a center position (O) in the flowing direction of the cooling air (A) on the wall surface of the cooling passage () between the adjacent turbulators ().

Blade cooling structure of gas turbine

Growing community of inventors

Takasago, Japan

Shunsuke Torii