A metal powder for powder metallurgy according to the invention contains Co as a principal component, Cr in a proportion of 25 to 32 mass %, Ni in a proportion of 5 to 15 mass %, Fe in a proportion of 0.5 to 2 mass %, W in a proportion of 4 to 10 mass %, Si in a proportion of 0.3 mass % to 1.5 mass %, and C in a proportion of 0.05 mass % to 0.8 mass %, wherein when one element selected from the group consisting of Ti, V, Y, Zr, Nb, Hf, and Ta is defined as a first element, and one element selected from the group and having a higher group number in the periodic table than that of the first element or having the same group number as that of the first element and a higher period number than that of the first element is defined as a second element, the first element is contained in a proportion of 0.01 to 0.5 mass %, and the second element is contained in a proportion of 0.01 to 0.5 mass %.

An inlet for an aircraft nacelle may comprise a nanoreinforced polyimide composite lip skin. The nanomaterials may increase thermal conductivity and decrease microcracking in the lip skin. A lip skin for an inlet with an electric heater may comprise a surface layer, an outer composite skin, an electric heater, an inner composite skin, and a thermal barrier coating. A lip skin for an inlet with a pneumatic deicing system may comprise a surface layer, a composite skin, and a thermal barrier coating.

A gas turbine arrangement, including a gas generator section (A), a power turbine section (B), and a generator section (C) coupled on a common shaft (10). The power turbine has its bearing block (12) provided with a copper cooling cup (9), which possesses a high thermal conductivity and conveys heat flux away from the side and block of the bearing and which has a design that enables the effect of a penetrating airflow.

The supercritical CO.sub.2 turbine in an embodiment includes: a rotary body; a stationary body housing the rotary body inside; and a turbine stage including a stator blade cascade in which a plurality of stator blades are supported inside the stationary body, and a rotor blade cascade in which a plurality of rotor blades are supported by the rotary body inside the stationary body, in which a supercritical CO.sub.2 working medium is introduced into the inside of the stationary body and flows via the turbine stage in an axial direction of the rotary body to thereby rotate the rotary body. Here, a thermal conductivity k1 and a specific heat c1 of a material constituting the rotary body and a thermal conductivity k2 and a specific heat c2 of a material constituting the stationary body satisfy a relationship represented by the following formula (A).

k1/c1≤k2/c2  formula (A)

High temperature stable thermal barrier coatings useful for substrates that form component parts of engines such as a component from a gas turbine engine exposed to high temperatures are provided. The thermal barrier coatings include a multiphase composite and/or a multilayer coating comprised of two or more phases with at least one phase providing a low thermal conductivity and at least one phase providing mechanical and erosion durability. Such low thermal conductivity phase can include a rare earth zirconate and such mechanical durability phase can include a rare earth a rare earth aluminate. The different phases are thermochemically compatible even at high temperatures above about 1200° C.

A turbocharger device includes a case having a turbine portion and a bearing portion connected to and extending from the turbine portion. The turbine portion defines a cavity that houses a turbine wheel and receives exhaust gas that rotates the turbine wheel. The bearing portion houses a shaft connected to the turbine wheel. The bearing portion has a radial thickness between an exterior surface and an interior surface. The interior surface defines a central channel. The bearing portion holds a bearing system that supports the shaft within the central channel. The bearing portion includes a lattice structure within the radial thickness. The lattice structure is a repeating three-dimensional array of frame segments connected to one another at junctions. The lattice structure engages a turbine back wall that is located between the turbine portion and the bearing portion. The lattice structure defines interstitial spaces between the frame segments.

A method for separating a first mechanical part from a second mechanical part is described, wherein the second mechanical part is bonded to the first mechanical part by an adhesive film along a connecting area, the first mechanical part having a first specific thermal conductivity and the second mechanical part having a second thermal conductivity that is higher than the first thermal conductivity. The method includes at least one cooling step during which the second mechanical part is cooled to a negative temperature and at least one stressing step during which the second mechanical part is subjected to mechanical stress in order to cause the adhesive film to break.

A method for separating a first mechanical part from a second mechanical part is described, wherein the second mechanical part is bonded to the first mechanical part by an adhesive film along a connecting area, the first mechanical part having a first specific thermal conductivity and the second mechanical part having a second thermal conductivity that is higher than the first thermal conductivity. The method includes at least one cooling step during which the second mechanical part is cooled to a negative temperature and at least one stressing step during which the second mechanical part is subjected to mechanical stress in order to cause the adhesive film to break.

A dual wall liner for a gas turbine engine may comprise a shell having a first side and a second side, a panel contacting the shell, the panel at least partially defining a hot gas path through which a hot gas flows, wherein the first side of the shell faces the panel, wherein the shell includes a thermal barrier coating (TBC) disposed on the first side of the shell. The TBC may thermally protect the shell from heat from a hot gas path.

A turbocharger (10A) has a back plate (41). The back plate (41) is provided with a plate section (41a) and a flange section (41c) which is formed radially outside the plate section (41a) and which is supported so as to be sandwiched between a bearing housing (16) and a turbine housing (31). The turbocharger (10A) is further provided with a flange heat shielding section (53) provided between the flange section (41c) and the bearing housing (16) and consisting of a material having lower heat conductivity than the turbine housing (31) and the back plate (41).