A composite fan casing for a gas turbine engine defining a central axis is generally provided. The composite fan casing includes a core having a plurality of core layers of reinforcing fibers bonded together with a thermosetting polymeric resin and having an outer surface. The composite fan casing further includes at least one stiffener integrally coupled to an aft portion of the outer surface of the core relative to the central axis. Additionally, the at least one stiffener comprises an elastic material.

A coating for a blade root/disk interface includes a layer of soft metal matrix, and a solid lubricant distributed through the soft metal matrix. Examples of materials include CuAl as the soft metal matrix and MoS.sub.2 as the solid lubricant, although others are also disclosed.

A rotor support system for a gas turbine engine may generally include a bearing assembly and a load reduction member configured to be coupled between the bearing assembly and a support frame of the engine. The load reduction member may include a fuse portion configured to fail when a load transmitted through the load reduction member exceeds a predetermined load threshold. The system may also include a load recoupling member provided between the bearing assembly and the support frame. The load recoupling member may be formed from a super-elastic shape memory alloy that allows the load recoupling member to undergo recoverable deformation without failing when the fuse portion fails such that the load recoupling member maintains a mechanical connection between the bearing assembly and the support frame.

A method of manufacturing a turbine vane within an engine case includes additively manufacturing a combustor liner within an engine case, additively manufacturing a support structure attached to the combustor liner at a radially distal position, and additively manufacturing the turbine vane attached to the support structure at an inwardly adjacent position to the radially distal position.

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.

Integrally bladed rotors (IBRs) are described. The IBRs include a central hub, an outer rim defining an outer circumference of the central hub, the outer rim defining a plurality of platforms, a plurality of circumferentially distributed blades, wherein a blade extends from each of the plurality of platforms, a rotor slot arranged between two adjacent blades, wherein the rotor slot is defined by a cut within the outer rim, and a rotor slot insert installed within the rotor slot, the rotor slot insert sized and shaped to fit within the rotor slot and prevent air leakage from a first side of the central hub to a second side of the central hub through the rotor slot during operation of the integrally bladed rotor.

A gas turbine engine has a compressor section including a lower pressure compressor and a higher pressure compressor, and a turbine section. A core engine housing surrounds the compressor section and the turbine section. An outer intermediate housing wall defines an internal chamber between the core housing and the outer intermediate housing. A fan rotor and a fan casing surround the fan rotor to define a bypass duct between the fan case and the outer intermediate housing. A heat exchanger is mounted in the internal chamber and receives high pressure air for cooling the high pressure air and delivering the high pressure air into the core engine housing to be utilized as cooling air for a component. Air from the lower pressure compressor is utilized to cool the higher pressure air in the heat exchanger.

A method of manufacturing a turbine vane within an engine case includes additively manufacturing a combustor liner within an engine case, additively manufacturing a support structure attached to the combustor liner at a radially distal position, and additively manufacturing the turbine vane attached to the support structure at an inwardly adjacent position to the radially distal position.

A blade for a gas turbine engine includes a body that includes an airfoil that extends in a radial direction from a 0% span position near an airfoil base to a 100% span position at an airfoil tip. The airfoil has a leading edge and a trailing edge that define the true chord length. The airfoil includes a first portion near the airfoil base with a first density and a second portion near the airfoil tip with a second density. The second density is less than the first density. The second portion includes an increasing true chord length in the radial direction. The second portion is in the range of 90% span to 100% span.

Turbine engine components are provided that have a repaired thermal barrier coating, along with their methods of formation and repair. The turbine engine component includes a thermal barrier coating on a first portion of a surface of a substrate; a repaired thermal barrier coating on a second portion of the surface of the substrate; and a ceramic coat on the outer bond coat. The thermal barrier coating includes an inner bonding layer and a first ceramic layer, with the inner bonding layer being positioned between the substrate and the first ceramic layer. The repaired thermal barrier coating generally includes an inner bond coat on the surface of the substrate and an outer bond coat on the inner bond coat. The inner bond coat is formed from a cobalt-containing material, while the outer bond coat is substantially free from cobalt.