An impeller for a rotary machine includes a base material of the impeller made of Al or an Al alloy. A surface layer for the impeller is formed by an electroless plating layer with a Ni—P based alloy and an under layer disposed between the base material and the surface layer, the under layer having a smaller Vickers hardness than the surface layer.

A snubber block for a manifold assembly of a gas turbine engine includes a body with at least one passage with a respective flow restrictive orifice, the body operable to receive a sensor and a male section that extends from the body, the male section receivable within a fuel manifold.

Using the systems and methods discussed herein, CMAS corrosion is inhibited via CMAS interception in an engine environment and/or is prevented or reduced by the formation of a metal oxide protective coating on a hot engine section component. The CMAS interception can occur while the engine is in operation in flight or in a testing or quality control environment. The metal oxide protective coating can be applied over other coatings, including Gd-zirconates (GZO) or yttria-stabilized zirconia (YSZ). The metal oxide protective coating is applied at original equipment manufacturers (OEM) and can also be applied in-situ using a gas injection system during engine use in-flight or during maintenance or quality testing. The metal oxide protective coating contains a rare earth element, aluminum, zirconium, chromium, or combinations thereof, and is from 1 nm to 3 microns in thickness.

A radial compressor impeller, having a hub-like main body, whose flow-guiding surface has a contour that increases in size in the radial direction from a flow inlet end in the direction of a flow outlet end, moving blades arranged on the hub-like main body, namely first moving blades and second moving blades. The first moving blades are formed longer than the second moving blades. Between each two adjacent first moving blades a second moving blade each is arranged. Third moving blades, which are formed shorter than the second moving blades are arranged between each adjacent first and second moving blade.

Embodiments of the present disclosure generally relate to protective coatings on turbine blades, turbine disks, and other aerospace components and methods for depositing the protective coatings. In one or more embodiments, a turbine blade includes a blade portion and a root coupled to the blade portion, where the root contains a protective coating disposed thereon. The protective coating is or contains one or more deposited crystalline film containing at least one of a metal oxide, a metal nitride, or a metal oxynitride and has a thickness of about 100 nm to about 10 μm. In some examples, a turbine blade assembly includes a disk and a plurality of the turbine blades coupled to the disk. The protective coating is disposed on the root on the turbine blade and/or a receiving surface on the turbine disk.

A blade comprises an airfoil extending from a trailing edge to a leading edge. The airfoil includes a body formed of an aluminum containing material. A sheath is at the leading edge and is formed of a titanium containing material. A sandwich is positioned intermediate the sheath and the airfoil body, the sandwich including an outer adhesive layer adjacent the sheath, an intermediate fabric layer and an inner adhesive layer adjacent the body. A gas turbine engine is also disclosed.

An exhaust gas turbine is provided. The exhaust gas turbine includes a first turbine housing part having insulating material extending along an interior surface and a second turbine housing part having insulating material extending along an interior surface, the second turbine housing part coupled to the first turbine housing part to form a volute directing exhaust gas to a turbine wheel.

A turbocharger impeller includes: a cylindrical boss portion disposed around a rotary axis; a hub portion connected to the boss portion and extends in a radial direction of the rotary axis; and a blade portion protruding from the boss portion and the hub portion toward a tip end side in the direction of the rotary axis and the radial direction. At least a part of an outer peripheral part of the hub portion in the radial direction or a part of the blade portion is provided with a resin second section and the second section is bonded to an aluminum first section. The turbocharger impeller includes the aluminum first section and the resin second section.

A method for manufacturing a turbine engine vane a root connected to a blade extending in a longitudinal direction includes the steps of providing a root; and providing mold with a first cavity and a second cavity that together define a recess in which the vane is formed. The recess includes a first space in which the blade is formed and a second space in which the root (1), c) is formed. The method further includes the steps of providing aluminum strips; positioning a fibrous reinforcement; arranging the vane root in the second space; and injecting a foam comprising aluminum or injecting an aluminum alloy into the first space of the recess of the mold such that the foam impregnates the fibrous reinforcement.

A method for manufacturing a blade for a turbine engine, including a root connected to a vane extending in a longitudinal direction, includes providing an assembly having a first part intended to form a root of the blade and a projecting second part projecting in the longitudinal direction from the first part; providing a mold comprising a first impression and a second impression delimiting together a cavity, said cavity comprising a first space and a second space; arranging the first part in the first space of the cavity and the second part in the second space of the cavity; and forming a third part by injecting an aluminium-based alloy in the cavity.