An exhaust-gas turbocharger (1) having a compressor housing (2); a turbine housing (3); and a bearing housing (4) which has a compressor-side flange (5) and which has a turbine-housing-side flange (6). The turbine-housing-side flange (6) is produced from a material which corresponds in terms of its mechanical and thermal properties to the material of the turbine housing (2).

A turbine exhaust case has an outer housing to be secured within a gas turbine engine and a central hub. Struts extend between the outer housing and the central hub. The struts are formed at least in part of a first material. The central hub is formed at least in part of a second material.

A turbine exhaust case has an outer housing to be secured within a gas turbine engine and a central hub. Struts extend between the outer housing and the central hub. The struts are formed at least in part of a first material. The central hub is formed at least in part of a second material.

The invention relates to a method for repairing a thermal barrier of a component comprising a substrate coated with such a thermal barrier, said substrate being made of a high-performance alloy, said thermal barrier being adhered to the alloy and having lower thermal conductivity than the alloy, the thermal barrier including at least one ceramic, one region of the thermal barrier being a region to be repaired, wherein said method includes the following steps: a) defining the region to be repaired, using a mask which protects the other regions of the thermal barrier; b) injecting a carrier gas loaded with droplets of ceramic precursor into a plasma discharge inside a plasma chamber of a plasma reactor containing the component to be repaired, while making the concentration of ceramic precursor in the carrier gas dependent on at least one parameter of the reactor selected from among: the pressure of the plasma chamber, the power of the plasma generator and the diameter of the precursor droplets, in order to control the stateliquid, gel or solidof the ceramic precursor having an effect on the region to be repaired; c) injecting a gas not loaded with ceramic precursor into a plasma discharge within the plasma chamber, steps b) and c) being repeated.

A component includes a substrate formed from a metallic or ceramic material and a thermal barrier coating positioned on the substrate. The component also includes a ceramic reflective coating integral with the thermal barrier coating. The reflective coating includes an arrangement of features configured to reflect at a wavelength at which peak emission from a heat source occurs. A method of making a component includes positioning a thermal barrier coating on the component and determining a wavelength emitted from a heat source. The method of making a component also includes producing an arrangement of features using a metamaterial to form a reflective coating and integrating the reflective coating with the thermal barrier coating.

The present invention provides a shaped component for a compressor wheel that can be used to manufacture a turbo compressor wheel which is excellent in terms of high-temperature strength, rigidity, and dynamic balance and, furthermore, has optimal performance suitable for diverse required characteristics in individual portions. In the shaped component for an aluminum alloy turbo compressor wheel of the present invention, a continuous casting rod-shaped material having a small diameter or a forging-completed material obtained by carrying out hot closed die forging on the continuous casting rod-shaped material having a small diameter is used as a material, and the continuous casting rod-shaped material or the forging-completed material has a structure in which the average numbers of grain boundaries intersected in a circumferential direction on a transverse section orthogonal to a casting direction and a forging direction are minimized in a central portion and are maximized in an outer circumferential portion, and the casting direction and the forging direction of the material are along a rotational center axis line direction of a compressor wheel.

A turbine component includes an airfoil-shaped core having an outer peripheral surface. A plurality of channels is formed in the core, each opening to the outer peripheral surface. The channels extend substantially radially to be elongated in the radial direction. A first platform is attached to the core. The component also includes an airfoil-shaped non-permeable skin having a hollow interior, an inner peripheral surface and an outer peripheral surface. The skin is sized so that the core can be received in the hollow interior of the skin. A second platform is attached to the skin. The core is received in the hollow interior of the skin such that the outer peripheral surface of the core engages the inner peripheral surface of the skin such that a plurality of generally radial cooling channels are formed between the channels in the core and the inner peripheral surface of the skin.

A gas turbine engine component includes a metal substrate and a coating system disposed on the metal substrate. The coating system includes at least one layer of aluminum-chromium oxide.

A ceramic component is generally provided that includes a silicon-based layer comprising a silicon-containing material (e.g., a silicon metal and/or a silicide) and about 0.001% to about 85% of a Ga-containing compound. For example, the silicon-based layer can be a bond coating directly on the surface of the substrate. Alternatively or additionally, the silicon-based layer can be an outer layer defining a surface of the substrate, with an environmental barrier coating on the surface of the substrate. Gas turbine engines are also generally provided that include such a ceramic component.

A method of making an aircraft blade is provided. The method comprises the steps of: assembling two or more fibre-reinforced thermoplastic composite parts into a blade assembly; and welding the fibre-reinforced thermoplastic composite parts together utilising an additional thermoplastic located at least at locations where the parts will abut when assembled. The additional thermoplastic has a melting or softening temperature lower than a melting temperature of each of the fibre-reinforced thermoplastic composite parts being assembled. The step of welding comprises heating the blade assembly to a temperature above the melting/softening temperature of the additional thermoplastic and below the melting temperature of each of the fibre-reinforced thermoplastic composite parts so as to melt/soften the additional thermoplastic and thereby weld the fibre-reinforced thermoplastic composite parts together to form the aircraft blade.