A turbine engine with an outer rotor that circumscribes an inner rotor or inner stator. The outer rotor includes circumferentially arranged components with a radial outer end and radial inner end. Inner ends of confronting sides of adjacent components include at least one damper element to dampen the relative motion of the components or to provide at least a partial seal between adjacent components.

A method of preventing spallation for a geometrically segmented thermally insulating top coat on an article, the method comprising forming a surface feature in a surface of the article; forming a crack deflection feature in the surface proximate the surface feature; disposing the thermally insulating topcoat over the surface feature; and forming segmented portions that are separated by faults extending through the thermally insulating topcoat from the surface feature.

The invention relates to a method for manufacturing a panel for supporting at least one cartridge of abradable material for a turbine engine casing, said panel including at least one block of honeycomb material covered with a rigid skin, characterised in that it includes: a step of machining (ET5) an outer surface of a block of material, in particular a honeycomb material, according to a three-dimensional profile configured to match that of an inner surface of the casing; and then, a step (ET6) of adhering the block of machined material to the rigid skin.

A filled abradable seal component, an associated method of manufacturing, and a turbomachine including the filled abradable seal component are disclosed. The method includes positioning the abradable seal component including a plurality of honeycomb cells, applying a filler material on the abradable seal component to fill the plurality of honeycomb cells, and curing the filler material at a temperature below 250 degrees Celsius to produce the filled abradable seal component. The filler material includes an abradable material, a binder material, and a fluid catalyst. The abradable material includes at least one of nickel chromium aluminum-bentonite, cobalt nickel chromium aluminum yttrium-polyester, cobalt nickel chromium aluminum yttrium-boron nitride, aluminum silicon-bentonite, aluminum bronze-polyester, nickel graphite, or aluminum silicon-boron nitride. The binder material includes at least one of aluminum, nickel-aluminum, aluminum thiophosphate, or aluminum thiosulfate. The fluid catalyst includes a solvent having hydroxyl groups.

Various embodiments include honeycomb structures including an abradable material, and a method of applying such honeycomb structures to steel components of a gas turbine engine in order to reduce rub damage. Particular embodiments include a honeycomb structure having a plurality of cells, each cell of the plurality of cells including a cell wall surrounding a void, and an abradable material within the void of each cell of the plurality of cells, the abradable material including a metallic alloy and hollow particles.

A fan case assembly for a gas turbine engine includes a honeycomb structure. The fan case assembly also includes a septum operatively coupled to the honeycomb structure and located radially inward of the honeycomb structure. The fan case assembly further includes a rubstrip in contact with the septum and located radially inward of the septum.

A method of forming a blade outer air seal includes applying a cavity layer that has a plurality of cavities to a radially inner side of a seal body. The cavity layer is coated with an abradable coating.

A cartridge for a fan case of a gas turbine engine includes an inlet acoustic liner section integrated with a thermally conforming liner section.

A sealing assembly (23) for a turbomachine (1) having a seal carrier (24) and a seal structure (30) configured on the seal carrier (24), the seal structure (30) having additively built-up projections (32) that extend in each case away from the seal carrier (24) to a free end (32.1), the projections (32) being constructed in each case to have a varying cross-sectional profile, namely a particular projection (32) in a section (33) that is distal to the seal carrier (24) having a smaller thickness (34) at the free end (32.1) than in a section (35) that is proximal to the seal carrier (24).

A seal assembly includes an annular carrier defining a radially inward face, and an opposed radially outward face defining a plurality of spline slots. The spline slots are configured and adapted to receive corresponding spline tabs to center the carrier in a surrounding structure. A honeycomb seal is mounted to the radially inward face of the carrier. The honeycomb seal is configured for sealing engagement with a knife-edge seal rotating relative to the honeycomb seal. The seal assembly has a relatively low coefficient of thermal expansion in comparison to the static component or the knife-edge seal. For example, a gas turbine engine can include a seal assembly as described above, wherein the carrier is mounted to a static component and a rotor mounted for rotation relative to the carrier, and wherein the rotor includes the knife edge seal component sealingly engaged to the honeycomb seal.