An integrally bladed rotor, including: a plurality of blades integrally formed with a hub as a single component, each of the plurality of blades having a blade body extending from the hub to an opposed blade tip surface along a longitudinal axis, wherein the blade body defines a pressure side and a suction side, and wherein the blade body includes a cutting edge defined between the blade tip surface of the blade body and the pressure side of the blade body, wherein the cutting edge is configured to abrade a seal section of an engine case. A method for manufacturing an integrally bladed rotor includes: forming a plurality of airfoils integrally with a hub to form a single component, each of the plurality of airfoils having an opposed tip surface with respect to the hub extending along a longitudinal axis, wherein each of the plurality of airfoils defines a pressure side and a suction side; and forming a cutting edge between the tip surface and the pressure side of each of the plurality of airfoils, wherein the cutting edge is configured to abrade a seal section of an engine case.

An impeller blade that includes an impeller blade body constructed of a first material. The impeller blade body defines a leading edge that faces a direction of rotation. A second material couples to the leading edge. The second material is a more erosion resistant material than the first material. The second material extends over the leading edge a distance to absorb high angle impacts of droplets and/or particulate. A third material couples to at least a portion of the impeller blade body.

The present invention discloses a turbine engine with at least a high pressure and a low pressure turbine section comprising a casing and at least one turbine blade rotatably mounted within the casing wherein at least part of the inner surface of the casing is covered with shrouds as abradables to provide clearance control between the inner surface and the tip of the at least one blade and wherein the tip of the blade is coated with a hard PVD coating, characterized in that the shroud material of at least the high pressure and/or the low pressure section comprises a porous ceramic based material and the hard PVD coating on the tip of the blade essentially consists of a droplet free nitride coating.

An article has a cavity defined by an inner surface, the cavity having a size such that a largest sphere placeable in the cavity has a diameter of less than 7 cm and a smallest sphere placeable in the cavity has a diameter of 0.5 mm; and a hard coating on the inner surface, the hard coating having a hardness between 18 to 100 GPa, the hard coating distributed on the inner surface such that a ratio of a coating thickness at a first region of the hard coating to that at a second region of the hard coating ranges from 0.75 to 1.33.

An anti-ice arrangement for a gas turbine engine may comprise an engine static structure, a fan blade housed for rotation within the engine static structure, and a magnetic field source mounted in close proximity to the fan blade and configured for inducing eddy currents in the fan blade to increase a surface temperature of the fan blade.

A heterogeneous composition is disclosed, including an alloy mixture and a ceramic additive. The alloy mixture includes a first alloy having a first melting point of at least a first threshold temperature, and a second alloy having a second melting point of less than a second threshold temperature. The second threshold temperature is lower than the first threshold temperature. The first alloy, the second alloy, and the ceramic additive are intermixed with one another as distinct phases. An article is disclosed including a first portion including a material composition, and a second portion including the heterogeneous composition. A method for forming the article is disclosing, including applying the second portion to the first portion.

Coating systems for a cooled turbine blade tip, such as a metal turbine blade tip, are provided. The coating system includes an abrasive layer overlying the surface of the turbine blade tip. One or more buffer layers may additionally be disposed between an outer surface of the blade tip and the abrasive layer. The coated blade tip can be used with a ceramic matrix composite (CMC) shroud coated with an environmental barrier coating (EBC) to provide improved cooling to the tip so as to lengthen oxidation time of the abrasive layer and reduce blade tip wear. Methods are also provided for forming the cooled blade tip and applying the coating system onto the cooled turbine blade tip.

A blade includes a blade body extending from a blade root to an opposed blade tip surface along a longitudinal axis. The blade body defines a pressure side and a suction side. The blade body includes a cutting edge defined where the tip surface of the blade body meets the pressure side of the blade body. The cutting edge is configured to abrade a seal section of an engine case. A method for manufacturing a blade includes forming an airfoil with a root and an opposed tip surface along a longitudinal axis, wherein the airfoil defines a pressure side and a suction side. The method also includes forming a cutting edge where the tip surface of the airfoil meets the pressure side of the airfoil.

An integrally bladed rotor, including: a plurality of blades integrally formed with a hub as a single component, each of the plurality of blades having a blade body extending from the hub to an opposed blade tip surface along a longitudinal axis, wherein the blade body defines a pressure side and a suction side, and wherein the blade body includes a cutting edge defined between the blade tip surface of the blade body and the pressure side of the blade body, wherein the cutting edge is configured to abrade a seal section of an engine case. A method for manufacturing an integrally bladed rotor includes: forming a plurality of airfoils integrally with a hub to form a single component, each of the plurality of airfoils having an opposed tip surface with respect to the hub extending along a longitudinal axis, wherein each of the plurality of airfoils defines a pressure side and a suction side; and forming a cutting edge between the tip surface and the pressure side of each of the plurality of airfoils, wherein the cutting edge is configured to abrade a seal section of an engine case.

A method for joining engine components includes positioning a first plurality of thermal protection structures across a thermal protection space between a first thermal protection surface and a second thermal protection surface. The first and second engine components are locally joined by forming a first plurality of transient liquid phase (TLP) or partial transient liquid phase (PTLP) bonds along corresponding ones of the first plurality of thermal protection structures between the first thermal protection surface and the second thermal protection surface. The second thermal protection surface is formed from a second surface material different from a first surface material of the first thermal protection surface.