Abstract: A gas atomization nozzle includes a through-hole formed along a center line; a nozzle portion configured of a Laval nozzle which is disposed around the center line and provided to be inclined at a predetermined angle toward the center line; and swirling motion imparting means for imparting a swirling flow around the center line to gas which is injected from the nozzle portion. The nozzle portion is formed in a ring shape which is continuous around the center line, and the swirling motion imparting means is configured as a fin provided in the nozzle portion to impart a swirling flow.

Abstract: Included in a gas turbine casing and a gas turbine are a first casing that has a first flange portion, first connecting holes, and first notches; a second casing that has a second flange portion, a plurality of second connecting holes, and second notches; and fastening bolts to fasten the first flange portion and the second flange portion with the portions being closely attached to each other and with each fastening bolt penetrating the corresponding first and second connecting holes. A first radial direction ratio (L/H) is set to be from 0.09 to 0.11 where H is a length of each of the first flange portion and the second flange portion in a radial direction, and L is a length of each of the first notch and the second notch in a radial direction.

Abstract: A metal member related to the present invention is provided with crystal grains of a metal and a granular reinforcing substance formed at boundaries of the crystal grains. The reinforcing substance includes grains of a shape with a grain area equivalent grain size larger than 1/100 of a grain area equivalent grain size of the crystal grains. The granular reinforcing substance preferably includes grains with a grain area equivalent grain size smaller than ? of the grain area equivalent grain size of the crystal grains. Additionally, the granular reinforcing substance preferably includes grains of a shape wherein a value of a length, in a first direction in which a length thereof is longest, divided by a length of a longest part in a direction orthogonal to the first direction is smaller than 5. A metal member with a high strength at high temperatures is manufactured by metal powder injection molding.

Abstract: A method for implementing an abradable coating comprising: a pattern-forming step in which, using a slurry containing ceramic particles and solvent, a slurry pattern is formed on the surface of a thermal barrier coating layer; and a firing step in which the slurry pattern formed on the surface of the thermal barrier coating layer is fired to form an abradable coating layer. A ceramic material included in the thermal barrier coating layer and ceramic particles included in the abradable coating layer are of the same type.

Abstract: 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.

Abstract: A virtual plane obtained by a locus when a tangent line, which passes through a tool distal end edge, of a virtual circle formed when a tool distal end edge of a cutting edge is rotated in a tool circumferential direction is moved parallel to a tool axial line is a reference plane, an angle (ARt) at which a cross line of the reference plane and the cutting edge virtual plane, is inclined with respect to the tool axial line projected onto the reference plane, is in a range of ?30 degrees to ?60 degrees, an angle (RR) at which a cutting edge tangent line which passes through the tool distal end edge and extends outward in a tool radial direction, is inclined with respect to a predetermined tool radial direction, which passes through the tool distal end edge, is in a range of ?30 degrees to ?75 degrees.

Abstract: A shaft portion and a blade portion provided on a side surface of the shaft portion are included, and the blade portion includes cutting blades arranged in a plurality of lines on a side surface of the shaft portion along a peripheral direction, and arranged in a plurality of stages in an extending direction of a shaft center of the shaft portion in each line. Further, the cutting blade has a radial-direction clearance angle, a tip end-side clearance angle, and a base end-side clearance angle.

Abstract: 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.

Abstract: An aircraft gas turbine is constituted by accommodating a compressor (14), a combustor (15), and a turbine (16) in a cylindrical main unit casing (12). A thick wall part (52) is provided on an outer periphery side of rotor blades (34) in the main unit casing. A cooling passage (53), for cooling the thick wall part (52) by circulating compressed air compressed by the compressor, is provided in the thick wall part. Also provided is a discharge passage (55) for discharging compressed air having circulated in the cooling passage to a combustion gas passage A. The structural strength of the thick wall part is ensured by appropriately cooling the thick wall part of the casing, while simplifying the structure and preventing a decrease in efficiency, thereby ensuring effective containment performance and an appropriate clearance between the casing and the rotor blades.

Abstract: A method of removing a work-affected layer formed on the worked surface of a TiAl-based alloy (base material) by machining work, without exerting any adverse effect on the base material. The method of removing the work-affected layer includes a step of dipping a TiAl-based alloy, having a work-affected layer formed on the surface thereof by machining, in an etchant containing predetermined concentrations of hydrofluoric acid and nitric acid. Within the etchant, the concentration of the hydrofluoric acid is not less than 5 g/L and not more than 56 g/L, and the concentration of the nitric acid is selected from within a range from not less than 50 g/L to not more than 260 g/L in accordance with a combination of the concentration of the hydrofluoric acid within the etchant and the etching treatment temperature.

Abstract: A gas turbine includes downstream side members that face a passage and have guide faces arranged in the passage side as it goes toward the upstream side. In a cross section including an axis center of a rotor shaft extending in a flow direction of a gas main stream, extension sections of the guide faces continuing to the upstream end portion of the guide faces are arranged downstream from the cavity inner wall portion of the upstream side member facing the cavities.