Hitachi, Ltd.

Mitsubishi Hitachi Power Systems, Ltd.

Central Research Institute of Electric Power Industry

Hitchi, Ltd.

Kunihiro Ichikawa

Shinya Imano

Hiroyuki Doi

Yoshitaka Kojima

Katsumi Tanaka

Yasushi Hayasaka

Yasuyuki Watanabe

Hiroyuki Endo

Shigeru Azuhata

Nobuyuki Iizuka

Nariyoshi Kobayashi

Takeshi Izumi

Takao Endo

Hideyuki Arikawa

Akira Mebata

Masaru Sekihara

Tadashi Kasuya

Hideaki Ishii

Yoshikazu Moritomo

Shigeo Sakurai

Tohru Hisamatsu

Tadashi Murakata

Yoshiharu Nakayama

Hiroshi Haruyama

Noriaki Matsuda

Takayuki Yoshioka

Nobuhiro Isobe

Takumi Konrai

Kazuto Mikazuki

Mitsutoshi Okada

Hidetoshi Kuroki

A gas turbine combustion chamber is provided, including a flow sleeve structure with an improved anti-vibration performance. A gas turbine combustion chamber of the present invention includes a liner, a transition piece, and a flow sleeve including a plurality of segments and integrated by welding a tie piece along joint portions of the segments. The tie piece includes a first member and a second member, the first member continuously extending along a longitudinal direction of the joint portions of the segments and being arranged to cover the joint portions, and the second member being formed at an end portion of the first member, having a width wider than the first member, and including a recess.

Gas turbine combustion chamber having a flow sleeve with a plurality of integrated segments

A gas turbine includes a compressor, a combustor, and a turbine. The turbine includes a honeycomb seal disposed so as to be secured to a casing side in a clearance between the casing and turbine blades rotating around a rotating shaft and a seal fin that is provided on an end face of each of the turbine blades facing the honeycomb seal. The seal fin extends in a direction perpendicular to the rotating shaft. The honeycomb seal is formed by a plurality of corrugated sheet metals overlapped with each other at walls of nodes thereof and the walls of the nodes are blazed with each other. Each of the corrugated sheet metals has trapezoids formed in alternating fashion. A longer direction of each wall of the nodes of the honeycomb seal is angled with respect to the rotational direction of the turbine blades.

Gas turbine with honeycomb seal

In a gas turbine blade where a part of the &#x3b3;' phase precipitation strengthened type Ni-based alloy base material is composed of a weld metal, the weld metal is a Ni-based alloy containing Ta from 4.8 to 5.3 wt. %, Cr from 18 to 23 wt. %, Co from 12 to 17 wt. %, W from 14 to 18 wt. %, C from 0.03 to 0.1 wt. %, Mo from 1 to 2 wt. %, and Al of 1 wt. % or less, in which the oxygen content is 0 to 30 ppm, the Ti content from 0 to 0.1 wt. %, and the Re content from 0 to 0.5 wt. %. A blade base metal is manufactured by the step of stripping, the step of solution heat treatment where the &#x3b3;&#x2032; phase is dissolved again, the step of welding in an inert gas chamber by a TIG method using a welding wire where the weld metal can be obtained, the step of HIP treatment at 1100&#xb0; C. to 1150&#xb0; C., and the step of an aging treatment at 835&#xb0; C. to 855&#xb0; C.

Gas turbine blade and manufacturing method thereof

A high-temperature resistant component for, e.g., a gas turbine hot part, includes an alloy substrate containing Ni, Co, or Fe as the principal component, and a thermal barrier coating formed over the surface of the substrate via a bond coat. The thermal barrier coating includes a porous thermal-barrier layer made of ceramic and an environmental barrier layer with corrosion resistance. An impregnated layer is provided between the environmental barrier layer and the thermal barrier layer. In the impregnated layer, the thermal barrier layer is impregnated with a part of the environmental barrier layer. The thermal barrier layer is made of a porous zirconia layer, and the environmental barrier layer includes silica as the principal component. The porous zirconia layer has pores impregnated with the part of the environmental barrier layer. As a result, the high-temperature resistant component has excellent corrosion resistance and excellent heat resistance.

High-temperature resistant component and gas turbine hot part

There is provided a high temperature component with a thermal barrier coating, which can be used as a high temperature component for a gas turbine, an aircraft gas turbine engine, or the like. A top coat is formed of a ceramic on a bond coat, the bond coat being formed on a heat resistant alloy substrate composed mainly of at least one element of nickel and cobalt, wherein the bond coat contains at least one of nickel and cobalt, chromium and aluminum, and further contains at least one selected from a group consisting of tantalum, cesium, tungsten, silicon, platinum, manganese and boron in a range of 0 to 20 wt %. The high temperature component according to the present invention has very high durability of a thermal-insulating ceramic layer, and is less susceptible to spalling damage.

High temperature component with thermal barrier coating and gas turbine using the same

A Ni-based alloy member has resistance against grain boundary fracture, fatigue strength, and oxidation resistance at temperatures near 1000&#xb0; C. or higher. The Ni-based alloy member includes a non-repaired region made of a Ni-based alloy base and a region repaired by welding, which is formed on the non-repaired region and which is made of a buildup-welded layer, the buildup-welded layer being made of a Ni-based alloy containing, by weight, 15% or less of Co, 18-22% of Cr, 0.8-2.0% of Al, 5.0% or less of Ta, 0.5% or less of Mo, 0.5% or less of Ti, 13-18% of W, 0.05-0.13% of C, 0.06% or less of Zr, 0.015% or less of B, 0.4-1.2% of Mn, and 0.1-0.3% of Si, the balance of the alloy being preferably essentially made of Ni.

Ni-based alloy member, method of producing the alloy member, turbine engine part, welding material, and method of producing the welding material

A thermal barrier coating comprises a bond coatingmade of high temperature corrosion resistance superior to a substrateand a thermal barrier coatingmade of ZrO2 series ceramics formed on the bond coating. The thermal barrier coating has cracks extending in the direction of the thickness of the barrier coat. Substantially all of the tips of the cracks have a distance between the tips and the boundary at the substrate side.

Heat resistant article having thermal barrier coating

A gas turbine rotor blade capable of effectively reducing creep damage by forming a cooling through hole to cool a target area in which significant creep damage of a shroud cover is predicted based on analysis of stress and temperature acting on the gas turbine rotor blade. A gas turbine using the rotor blade, and a power plant using the gas turbine are also provided. The gas turbine rotor blade includes a blade section provided with a shroud cover at an outer peripheral end thereof, and a platform, a shank and a dovetail which are formed in integral structure to successively continue from the blade section. An inner cooling hole is formed to penetrate through the gas turbine rotor blade from the dovetail to the shroud cover. The shroud cover has a cooling through hole formed to open in an outer surface of the shroud cover and extend in communication with the inner cooling hole.

Gas turbine rotor blade, gas turbine using the rotor blade, and power plant using the gas turbine

A welding material composition, which is a nickel based super alloy having &#x3b3;' phase and chromium carbides precipitated. The composition comprising 18 to 25% by weight of Co, 15 to 20% by weight of Cr, 1.5 to 5.5% by weight of Al, 5 to 14% by weight of W, 0.05 to 0.15% by weight of C, 0 to 0.02% by weight of B, 0 to 1% by weight of at least one of Ta, Nb, Ti, Mo, Re and Fe, 0 to 0.5% by weight of at least one of V, Zr, rare earth elements and Y, 0 to 1% by weight of Mn, 0 to 0.5% by weight of Si, and the balance being Ni.

Welding material, gas turbine blade or nozzle and a method of repairing a gas turbine blade or nozzle

A life management system for high-temperature parts of a gas turbine has one server system and a plurality of client systems , and , all of which are connected via an Intranet. Further, the server system manages a program for performing evaluation of the remaining life and the life management, and each of the client systems , and has a subprogram for accessing the database and for entering data respectively. The clients are dedicated to different objects and share respective element data such as real component damage, design, materials, etc. which are necessary for the evaluation of the remaining life. Further, this system enables the operation of the gas turbine to be optimized based on the damage of the evaluated parts, hence contributing to operational cost reduction of the gas turbine.

Service life management system for high-temperature part of gas turbine

A gas turbine combustor is provides with a stabilizer at a downstream side of a premixing device for premixing and supplying fuel and air into a combustion chamber. The stabilizer has a cylindrical part extending axially from an upstream end thereof to a downstream end and a stabilizing part at the downstream end of the cylindrical part for stabilizing a flame, and the stabilizer is mounted on an inside of the combustor by a plurality of members arranged annularly and secured to the inside of the combustor at their ends, each of the members having a part slidably inserted in an opening of the cylindrical part so as to restrict the stabilizer to move axially while allowing deformation of the stabilizer and another part supporting the upstream end of the cylindrical part so as to restrict radial movement of the stabilizer.

Gas turbine combustor and gas turbine

Growing community of inventors

Hitachinaka, Japan

Kunihiro Ichikawa