Abstract: An aerospace component, for example, used in a gas turbine engine, includes the following structurally-integrated layers: a metallic layer and a composite layer having reinforcing fibers embedded in a matrix material. The aerospace component can also include an insulating layer disposed between the metallic layer and the composite layer where the insulating layer has a thermal conductivity that is lower than a thermal conductivity of the composite layer.

Abstract: A composite case for a gas turbine engine includes an inner structure assembly defining an impact containment zone. The inner structure assembly is configured to circumscribe a plurality of rotatable blades of the gas turbine engine and to receive blade fragments in an event of blade failure. The inner structure assembly includes one or more layers containing a resin impregnated therein, and an outer structure assembly surrounding the inner structure assembly and integrally molded thereon. The outer structure assembly includes one or more layers containing a resin impregnated therein. One or more layers of the inner and the outer structure assemblies include resin reinforced by nanoparticles.

Abstract: A hybrid vane for a gas turbine engine. The hybrid vane comprises an airfoil having an inner core composed of a fiber-reinforced thermoplastic composite. A longitudinal axis of the hybrid vane extends between a vane root and a vane tip. The hybrid vane further comprises a metallic outer layer at least partially covering the inner core.

Abstract: A fan blade for a fan of a gas turbine engine is described which includes an airfoil having an inner core and an outer shell composed of different metals, and a galvanic separator therebetween. The galvanic separator including an adhesive layer covering said at least a portion of the inner core, and a non-conductive fabric covering the adhesive layer. A plurality of solid metal particles may be disposed on an outer surface of the non-conductive fabric layer, between the non-conductive fabric layer and the outer shell.

Abstract: A gaspath traversing component of a gas turbine engine comprises a wall having an outer edge surface and a thickness relative to the gaspath, the wall having a plurality of layers of composite materials forming the thickness. A wear indication layer is embedded within the plurality of layers of composite material, the wear indication layer being visually contrasting with the composite material. The wear indication layer is positioned interiorly of at least one layer of said plurality of layers of composite material relative to the outer edge surface. A method for attending to a gas traversing component of a gas turbine engine is also provided.

Abstract: An aerospace component, for example, used in a gas turbine engine, includes the following structurally-integrated layers: a metallic layer and a composite layer having reinforcing fibers embedded in a matrix material. The aerospace component can also include an insulating layer disposed between the metallic layer and the composite layer where the insulating layer has a thermal conductivity that is lower than a thermal conductivity of the composite layer.

Abstract: A method of applying a nanocrystalline metal coating to a titanium or titanium alloy blade of a gas turbine engine is described. The method includes selecting an intermediate bond coat from the group consisting of nickel, nickel alloy, P, B, Tl and combinations thereof, and applying the intermediate bond coat to at least a portion of the blade. A nanocrystalline metal coating having an average grain size of between 10 nm and 500 nm is then applied to the portion of the blade overtop of the intermediate bond coat.

Abstract: A composite case for a gas turbine engine includes an inner structure assembly defining an impact containment zone. The inner structure assembly is configured to circumscribe a plurality of rotatable blades of the gas turbine engine and to receive blade fragments in an event of blade failure. The inner structure assembly includes one or more layers containing a resin impregnated therein, and an outer structure assembly surrounding the inner structure assembly and integrally molded thereon. The outer structure assembly includes one or more layers containing a resin impregnated therein. One or more layers of the inner and the outer structure assemblies include resin reinforced by nanoparticles.

Abstract: A fan blade for a fan of a gas turbine engine is described which includes an airfoil having an inner core and an outer shell composed of different metals, and a galvanic separator therebetween. The galvanic separator including an adhesive layer covering said at least a portion of the inner core, and a non-conductive fabric covering the adhesive layer. A plurality of solid metal particles may be disposed on an outer surface of the non-conductive fabric layer, between the non-conductive fabric layer and the outer shell.

Abstract: Sensor probes with anti-icing and methods for manufacturing such sensor probes are disclosed. An exemplary sensor probe comprises a probe sensor, a probe body housing the probe sensor and an anti-icing device. The anti-icing device is bonded to and covers at least part of the probe body for preventing ice from forming on the probe body. The anti-icing device comprises a heating element thermally insulated from the probe body by a thermal insulator and a hydrophobic coating covering the heating element and defining an outer surface for exposure to an airflow. The heating element may be deposited using a direct write process.

Abstract: An inner bypass duct of a gas turbine engine includes a plurality of non-structural front panels arranged circumferentially. Each front panel has an outer surface and an inner surface. The inner surface is at least partially covered by a composite structure. The composite structure includes a fireproof layer and an acoustic layer disposed between the fireproof layer and the inner surface. A non-structural panel for an inner bypass duct and a method of forming such panel are also presented.

Abstract: An acoustic structure for a gas turbine engine comprising a noise reduction layer and a fire protector layer connected to the noise reduction layer. The noise reduction includes a perforated inner wall adapted to be in contact with a first fluidic environment, and a noise reduction adjacent to the inner wall. The fire protector layer includes a non-perforated outer wall adapted to be in contact with a second fluidic environment having potentially a fire, a fire protector adjacent to the outer wall, and a pressure resisting wall disposed between the fire protector and the noise reduction. The second fluidic environment is under a pressure lower than a pressure of the first fluidic environment. The inner and outer walls are load-bearing walls of the acoustic structure.

Abstract: A component of a gas turbine engine comprises a substrate, a corrosion resistant top layer, and an intermediate corrodible layer disposed between the corrosion resistant top layer and the substrate. When corroding, the intermediate layer has a color contrasting with a color of the top layer. A method of detecting a crack when it penetrated the top layer in a component of a gas turbine engine having a corrosion resistant top layer and an intermediate corrodible layer comprises, in sequence, observing that at least one area of the component has a color contrasting with that of the top layer; determining that the color of the at least one area is a result of corrosion of the intermediate corrodible layer; and determining that the top layer has a crack as a result of determining corrosion of the intermediate layer. A method of facilitating crack detection in a component is also presented.

Abstract: A component of a gas turbine engine comprises a substrate, a corrosion resistant top layer, and an intermediate corrodible layer disposed between the corrosion resistant top layer and the substrate. When corroding, the intermediate layer has a color contrasting with a color of the top layer. A method of detecting a crack when it penetrated the top layer in a component of a gas turbine engine having a corrosion resistant top layer and an intermediate corrodible layer comprises, in sequence, observing that at least one area of the component has a color contrasting with that of the top layer; determining that the color of the at least one area is a result of corrosion of the intermediate corrodible layer; and determining that the top layer has a crack as a result of determining corrosion of the intermediate layer. A method of facilitating crack detection in a component is also presented.

Abstract: Sensor probes with anti-icing and methods for manufacturing such sensor probes are disclosed. An exemplary sensor probe comprises a probe sensor, a probe body housing the probe sensor and an anti-icing device. The anti-icing device is bonded to and covers at least part of the probe body for preventing ice from forming on the probe body. The anti-icing device comprises a heating element thermally insulated from the probe body by a thermal insulator and a hydrophobic coating covering the heating element and defining an outer surface for exposure to an airflow. The heating element may be deposited using a direct write process.

Abstract: A method of manufacturing a vane component of a gas turbine engine is described, and includes forming a core of the vane component using a rapid manufacturing process, and applying a coating of nanocrystalline metal onto the core, the nanocrystalline metal fully encapsulating the core and defining an outer structural surface of the vane component.

Abstract: An inner bypass duct of a gas turbine engine includes a plurality of non-structural front panels arranged circumferentially. Each front panel has an outer surface and an inner surface. The inner surface is at least partially covered by a composite structure. The composite structure includes a fireproof layer and an acoustic layer disposed between the fireproof layer and the inner surface. A non-structural panel for an inner bypass duct and a method of forming such panel are also presented.

Abstract: A process for forming an abradable coating on a component of a gas turbine engine includes providing a particulate mixture including an abradable coating material in particle form, the abradable material including at least a solid lubricant, and applying the particulate mixture containing the abradable coating material onto the component using pulsed shockwaves carrying the particular mixture, while maintaining a temperature of the particular mixture below a threshold temperature above which the solid lubricant in the abradable material substantially reacts chemically.

Abstract: A component of a gas turbine engine comprises a substrate, a corrosion resistant top layer, and an intermediate corrodible layer disposed between the corrosion resistant top layer and the substrate. When corroding, the intermediate layer has a color contrasting with a color of the top layer. A method of detecting a crack when it penetrated the top layer in a component of a gas turbine engine having a corrosion resistant top layer and an intermediate corrodible layer comprises, in sequence, observing that at least one area of the component has a color contrasting with that of the top layer; determining that the color of the at least one area is a result of corrosion of the intermediate corrodible layer; and determining that the top layer has a crack as a result of determining corrosion of the intermediate layer. A method of facilitating crack detection in a component is also presented.

Abstract: A method of forming an abradable coating on a gas turbine engine component comprising, in sequence: placing dry lubricant particles and trapping particles in a channel having a spraying end and containing a gas; causing at least one shockwave in the gas to travel in the channel toward the spraying end, the at least one shockwave causing the dry lubricant particles and the trapping particles to travel in the channel with it, the at least one shockwave reducing interparticle spacing and increasing particles density; directing a resulting flow of the dry lubricant particles and the trapping particles from the spraying end at a supersonic velocity to impact the component; and then plastically deforming the trapping particles upon impacting the component with the resulting flow thereby trapping the dry lubricant particles with the deformed trapping particles onto the component to provide the abradable coating.