Mitsubishi Heavy Industries Aero Engines, Ltd.

Osaka Yakin Kogyo Co., Ltd.

Mitsubishi Heavy Industries Limited

Tadayuki Hanada

Kenji Suzuki

Takayuki Kurimura

Mineaki Matsumoto

Hisashi Kitagaki

Yohei Fujimoto

Yasuyuki Uetsuki

Kazuki Hanami

Takafumi Ota

Kosuke Nishikawa

Takaaki Hase

Azusa Tamugi

Shuntaro Terauchi

Tomoyuki Onishi

Hirokazu Hagiwara

Daigo Fujimura

Osamu Morii

Syuntaro Terauchi

Hiroshi Watanabe

Yoshihiro Kuwamura

Kentaro Shindo

Eiichi Yanagisawa

Mitsuru Kondo

Yoshiyuki Okabe

Shigenari Horie

Ryutaro Fujisawa

Hiroji Tada

Shintaro Sobu

Satoru Yamazaki

Tomohiro Imae

Kunihiro Shimizu

Kenji Doi

Yusuke Ichihashi

By melting a shaping material in which a metal powder and a binder are mixed and by carrying out injection molding (primary shaping) in an injection mold, an injection molded body, or an intermediate shaped body are produced. The injection molded body or the intermediate shaped body is placed by a transfer mold and is subjected to a gravity shaping (secondary shaping) with a transformation. A sintered body is manufactured by carrying out debindering and sintering to the injection molded body.

Sintered body, method of manufacturing sintered body, combustor panel, and method of manufacturing combustor panel

Provided are a coated member in which damage of a coating film can be suppressed in a high temperature environment and the coating may be performed at low cost, and a method of manufacturing the same. A coated member includes a bond coat and a top coat sequentially laminated on a substrate made of a Si-based ceramic or a SiC fiber-reinforced SiC matrix composite, wherein the top coat includes a layer composed of a mixed phase of a (YLn)SiOsolid solution (here, Lnis any one of Nd, Sm, Eu, and Gd) and YSiOor a (YLn')SiOsolid solution (here, Ln&#x2032; is any one of Nd, Sm, Eu, and Gd), or a mixed phase of a (YLn)SiOsolid solution (here, Lnis any one of Sc, Yb, and Lu) and YSiOor a (YLn&#x2032;)SiOsolid solution (here, Ln&#x2032; is any one of Sc, Yb, and Lu).

Coated member and method of manufacturing the same

A hydrogenation-dehydrogenation method for a TiAl alloy includes performing hydrogenation treatment of the TiAl alloy in an environment of a set temperature equal to or higher than a temperature at which phase transformation to a &#x3b2; phase starts; and performing dehydrogenation treatment of the TiAl alloy which has been subjected to the hydrogenation treatment. The set temperature ranges from 1,100&#xb0; C. to 1,600&#xb0; C.

Hydrogenation-dehydrogenation method for TiAl alloy and method for producing TiAl alloy powder

A coating method includes depositing a slurry including a coarsely particulate ceramic and a finely particulate ceramic on a base material configured with an oxide-based ceramics matrix composite such that a proportion of coarse particles decreases towards a surface of the base material; forming a bond coating by performing a heat treatment on the base material on which the slurry has been deposited; and forming a top coating by thermally spraying a ceramic onto the bond coating. The oxide-based ceramics matrix composite is an alumina silica type oxide-based ceramics matrix composite. The coarsely particulate ceramic and the finely particulate ceramic are alumina-based powder.

Coating method, coating layer, and turbine shroud

An aircraft component is used for an aircraft gas-turbine engine. The aircraft component includes an annular part, a flange, and a boss. The annular part has an outer circumferential surface. The flange is formed at one end portion of the annular part in an axial direction. The boss projects from the outer circumferential surface of the annular part to the radial direction. On a section cut along an axial direction of the annular part, the outer circumferential surface of the annular part between the flange and the boss has a taper part that is formed into a tapered shape in which plate thickness becomes thicker from the flange toward the boss.

Aircraft component and aircraft gas-turbine engine

Provided are a coated member in which damage of a coating film can be suppressed in a high temperature environment and the coating may be performed at low cost, and a method of manufacturing the same. A coated member includes a bond coat and a top coat sequentially laminated on a substrate made of a Si-based ceramic or a SiC fiber-reinforced SiC matrix composite, wherein the top coat includes a layer composed of a mixed phase of a (YLn)SiOsolid solution (here, Lnis any one of Nd, Sm, Eu, and Gd) and YSiOor a (YLn)SiOsolid solution (here, Lnis any one of Nd, Sm, Eu, and Gd), or a mixed phase of a (YLn)SiOsolid solution (here, Lnis any one of Sc, Yb, and Lu) and YSiOor a (YLn)SiOsolid solution (here, Lnis any one of Sc, Yb, and Lu).

The present invention enables machining to be easily performed and reduces stress concentration on a machined hole that is formed in a cylindrical member. Recesses () recessed in the depth direction of a machined hole () are formed on circumferential side-portions of the machined holeformed in a cylindrical member (). In each of the recesses (), a part of the opening edge is formed to be a circular arc portion () that has a circular arc shape, the bottom is formed to be gradually shallowed by an inclined surface () toward an opened portion () in which the circular arc shape is opened, from a part along the circular arc portion (), and the circular arc portion () is disposed toward the machined hole ().

Stress reduction structure, gas turbine casing, and gas turbine

Provided is a transition duct that forms an annular gas flow channel through which a main flow gas flows from a high-pressure turbine to a low-pressure turbine, wherein the gas flow channel has an inner peripheral flow channel surface and an outer peripheral flow channel surface, the inner peripheral flow channel surface and the outer peripheral flow channel surface extend radially outward while angles of inclination relative to an axial direction of a rotating shaft change from the high-pressure turbine (first turbine) toward the low-pressure turbine (second turbine), and an inner peripheral maximum inclined part is provided in a range Athat extends in the axial direction from a position of alignment with an outer peripheral maximum inclined part to a position advanced toward the low-pressure turbine by a length no more than 20% of the duct length.

Transition duct, turbine, and gas turbine engine

A gas turbine for an aircraft includes: blades each including a blade body extending outward from a rotor rotating about an axis in a radial direction of the rotor and a fin protruding from an outer peripheral surface of a blade shroud formed on a distal end of the blade body; a casing installed to form gaps between the casing and the fins; and vanes each including a vane shroud formed such that at least a leading end portion of the vane shroud formed upstream in an axial direction is located outside with respect to an extended line of an inner peripheral surface of the casing in the radial direction.

Gas turbine for aircraft

An aircraft component is used in a gas turbine engine for an aircraft. The aircraft component includes an annular part having an outer peripheral surface, and a boss part protruding from the outer peripheral surface of the annular part in a radial direction. In the boss part, at least two through-holes are formed to penetrate the boss part in the radial direction at predetermined intervals. In the boss part around the two through-holes, a cut-out part where a part of the boss part is cut out is formed.

Aircraft component and gas turbine engine for aircraft

A coated member includes a heat-shielding coating layer made of a zirconia-dispersed silicate in which ytterbia-stabilized zirconia is precipitated as a dispersed phase in a matrix phase which is any one of a rare earth disilicate, a rare earth monosilicate, and a mixed phase of the rare earth disilicate and the rare earth monosilicate. The rare earth disilicate is a (Y[Ln])SiOsolid solution wherein Lnis any one of Sc, Yb, and Lu, or a (Y[Ln])SiOsolid solution wherein Lnis any one of Nd, Sm, Eu, and Gd. The rare earth monosilicate is YSiO, [Ln']SiO, a (Y[Ln&#x2032;]b)SiOsolid solution wherein Ln&#x2032; is any one of Sc, Yb, and Lu, or a (Y[Ln&#x2032;])SiOsolid solution wherein Ln&#x2032; is any one of Nd, Sm, Eu, and Gd.

Coated member, coating material, and method of manufacturing coated member

A method for metal powder injection molding includes injecting a first metal powder of a TiAl-based intermetallic compound into a mold, and molding the first metal powder through use of an injection molding machine; injecting a second metal powder of a TiAl-based intermetallic compound having a same constituent as the first metal powder and having a different average particle diameter from the first metal powder into a mold, and molding the second metal powder through use of the injection molding machine; and sintering molded articles obtained by molding the first metal powder and the second metal powder, and producing a mixed sintered compact in which a first sintered compact of the molded article obtained by molding the first metal powder and a second sintered compact of the molded article obtained by molding the second metal powder are integrated.

Method for metal powder injection molding

A gas turbine includes a compressor, a combustor, and a turbine. The compressor, the combustor, and the turbine are arranged along an extending direction of a rotating shaft. The gas turbine also includes a combustor casing housing the combustor and a turbine casing housing the turbine. The combustor casing and the turbine casing are joined to each other via respective flanges thereof projecting toward the outside. The gas turbine further includes a projection part on an inner face of the combustor casing. The projection part projects toward the inside in a radial direction in at least part of a range in the extending direction of the rotating shaft between the flanges and an end of a combustion chamber on the compressor side in the combustor.

Gas turbine combustor casing having a projection part

A method for constructing an abradable coating comprises: a slurry layer formation step Sin which a slurry layeris formed on the surface of a base materialusing a slurry containing ceramic particles and a solvent; a calcination step Sin which the slurry layerformed on the surface of the base materialis sintered and a sintered layerto be a portion of an abradable coating layeris formed; and a slurry removal step Sin which extraneous slurry is removed after the abradable coating layerhas been formed on the surface of the base material, a plurality of the sintered layershaving been laminated in the abradable coating layerthrough a plurality of repeated cycles of the slurry layer formation step Sand the calcination step S

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Aichi, Japan

Tadayuki Hanada