An austenitic stainless steel alloy and turbocharger kinematic components are provided. An austenitic stainless steel alloy includes, by weight, about 23% to about 27% chromium, about 18% to about 22% nickel, about 0.5% to about 2.0% manganese, about 1.2% to about 1.4% carbon, about 1.6% to about 1.8% silicon, about 0.2% to about 0.4% nitrogen, about 0% to about 0.5% molybdenum, sulfur in an amount of less than about 0.01%, phosphorous in an amount of less than about 0.04%, and a balance of iron, and other inevitable/unavoidable impurities that are present in trace amounts. The turbocharger kinematic components are made at least in part using this stainless steel alloy.

An improved stator assembly for use in a gas-turbine engine is disclosed. The stator assembly may comprise a vane, an inner diameter (ID) ring, an outer diameter (OD) ring, a vane disposed between the ID ring and the OD ring, a potting component coupling the vane to at least one of the OD ring or the ID ring, and a potting embedded component disposed within the potting component. The potting embedded component may prevent disbond of the potting component during operation of the gas-turbine engine.

Particulate composite materials and devices comprising the same are provided.

The present disclosure provides a method for manufacturing a preform made of reinforcement fibers woven in a longitudinal direction. The preform is impregnated with resin in order to form an elongated element having a variable transverse cross-section. In one form, the method includes the step of simultaneously including a reduction (or increase) in width and an increase (or reduction) in height. The variable cross-section includes, along the length thereof, a consistent number (c) of continuous warp threads arranged in layers. The method for reducing width includes carrying out a change in weave, adding additional weft threads, and simultaneously drawing the teeth of the longitudinal beater reed closer together so as to increase the number of layers.

A turbine nozzle for a turbine engine, including two annular walls extending about the same axis, the walls being connected to each other by blades having an aerodynamic profile, the nozzle being divided into sectors and including several nozzle sectors arranged circumferentially end-to-end about a longitudinal axis. The blades include first blades made from a first metal alloy, and second blades made from a second metal alloy different from the first alloy, each nozzle sector including first blades and second blades. The invention also relates to a turbine engine, in particular of an aircraft, including at least one such nozzle. The invention also relates to a method for manufacturing such a nozzle in which each of the sectors is produced, for example, by additive manufacturing.

A turbine system including a sealing component is presented. The sealing component includes a ceramic material. The ceramic material includes grains having an average grain size of less than 10 microns. A turbine shroud assembly including the sealing component is also presented.

The present disclosure is directed to an engine component for a gas turbine engine, the engine component including a substrate that includes a composite fiber and defines a surface. An energy harvesting fiber is positioned within the substrate.

Provided is a turbocharger with further improved thermal efficiency. The turbocharger includes a heat shielding plate disposed between a center housing and a turbine housing. The center housing rotatably and pivotally supports a rotating shaft that connects a turbine impeller and a compressor impeller, and the turbine housing houses the turbine impeller. The turbocharger includes: a first flange portion; a second flange portion formed corresponding to the first flange portion; a clamp member combining the first flange portion and the second flange portion and fixing the mutual positional relationship; and a heat insulating ring interposed between an end surface, which is a heat shielding plate fixing portion that fixes the heat shielding plate, and the heat shielding plate on an inner peripheral side with respect to the first flange portion and the second flange portion.

A method for forming a composite part of a gas turbine engine. The method includes assembling the composite part of a first composite material and a second composite material. The second composite material defines an outer surface of the composite part, and is selected to be curable at a cure temperature generated by heat from operation of the engine. The first composite material is selected to have an operating temperature limit less than the cure temperature. The method includes placing the composite part within the engine so that, in use, the second composite material is cured by exposure to the heat generated from operation of the engine. The second composite material thermally shields the first composite material from the heat generated from operation of the engine. The method includes operating the engine to cure the second composite material.

Yttria stabilized zirconia (YSZ) particles (40) form a thermal barrier layer (58) on a metal substrate (24). The YSZ particles have a porous interior (52, 54) and a fully melted and solidified outer shell (50). The thermal barrier layer may have porosity greater than 12%, including porosity within the particles and inter-particle gap porosity. Inter-particle gaps may be greater than 5 microns. The thermal barrier layer may exhibit elastic hysteresis and an average modulus of elasticity of 15-25 GPa. A bond coat (44A, 44B) may be applied between the substrate and the thermal barrier layer. The bond coat may have a first dense MCrAlY layer (44A) on the substrate and a second rough, porous MCrAlY layer (44B) on the first MCrAlY layer, the bond layers diffusion bonded to each other and to the substrate.