Abstract: A compressor with negative coefficient of thermal expansion case material comprising a rotor having blades with tips, the case including an inner case comprising a negative coefficient of thermal expansion material, and a tip clearance located between the tips and the inner case; wherein the tip clearance is maintained responsive to a flow of air over the negative coefficient of thermal expansion material.

Abstract: A combustor liner panel attachment assembly includes a liner extending from a first end to a second end, and circumferentially to partially define a combustion zone. The assembly also includes a spring element located adjacent to a portion of the liner and operatively coupled to a stationary structure, the spring element having a recessed segment. The assembly further includes a protrusion feature extending radially outwardly from the liner, the protrusion feature disposed within the recessed segment of the spring element to axially retain the liner.

Abstract: A gas turbine engine includes a main compressor section having a downstream most location, and a turbine section, with both the main compressor section and the turbine section housing rotatable components. A first tap taps air compressed by the main compressor section at an upstream location upstream of the downstream most location. The first tap passes through a heat exchanger, and to a cooling compressor. Air downstream of the cooling compressor is selectively connected to reach at least one of the rotatable components. The cooling compressor is connected to rotate at a speed proportional to a rotational speed in one of the main compressor section and the turbine section. A valve system includes a check valve for selectively blocking flow downstream of the cooling compressor from reaching the at least one rotatable component. A dump valve selectively dumps air downstream of the cooling compressor. A method is also disclosed.

Abstract: Aspects of the disclosure are directed to a seal comprising: a shoe, and at least two beams coupled to the shoe, where a first width associated with the beams exceeds a second width associated with the shoe in a reference direction. Aspects of the disclosure are directed to an engine comprising: a compressor section, a turbine section, and a floating, non-contact seal that includes: a shoe, and at least two beams coupled to the shoe, where the beams extend past an edge of a flowpath surface associated with the shoe in a reference direction.

Abstract: A piston seal assembly for a gas turbine engine includes a seal composed of a nickel-based superalloy; a component in contact with the seal and defining a seal-counterface; and a coating on the seal at the seal-counterface, wherein the coating is a metal alloy binder phase and a hard particle phase distributed through the binder phase.

Abstract: A variable vane actuation system for a gas turbine engine includes a stem section that forms a base and a contoured section that extends from the base along an axis. A vane arm comprising a claw section received onto the contoured section and a fastener fastened to the contoured section to load the claw section to the base.

Abstract: An improved face seal may comprise a seal housing having an annular extrusion, and a primary seal coupled to the seal housing. The primary seal comprises a sealing face and a base opposite the sealing face, wherein the primary seal has an annular structure coaxial with the annular extrusion with the base proximate a first end of the seal housing. The sealing face comprises a disrupted surface.

Abstract: A blade outer air seal assembly includes a carrier that has a slot and a hole that extends into the slot. A blade outer air seal has a plurality of segments that extend circumferentially about an axis and mounted in the carrier. At least one of the plurality of segments has a base portion and a first wall that extends axially and radially outwardly from the base portion. The first wall has an aperture. A pin extends through the hole and the aperture.

Abstract: A method for forming an ultra-high temperature (UHT) composite structure includes dispensing a polymeric precursor with a spinneret biased at a first DC voltage; forming a plurality of nanofibers from the polymeric precursor; receiving the plurality of nanofibers with a collector biased at a second DC voltage different than the first DC voltage; and changing a direction of movement of the plurality of nanofibers between the spinneret and the collector with a plurality of magnets having a magnetic field by adjusting the magnetic field.

Abstract: A vane arc segment includes an airfoil piece that has a first platform, a second platform, and an airfoil section between the first platform and the second platform. At least the first platform is formed of a fiber-reinforced composite (FRC) and defines a radial flange. A cap is fitted on the radial flange.

Abstract: A blade outer air seal assembly includes a support structure. A blade outer air seal has a plurality of segments that extends circumferentially about an axis and is mounted in the support structure. At least two of the segments have a base portion that extends from a first circumferential side to a second circumferential side. A first protrusion extends from the first circumferential side and has a first radially extending slot. A second protrusion extends from a second circumferential side and has a second radially extending slot. A feather seal is arranged in the first radially extending slot and the second radially extending slot between at least two segments.

Abstract: A method of designing a gas turbine engine includes locating purge openings in fluid communication with a first stage cavity. At least one of a cover plate or a rotor disk is positioned adjacent the first stage cavity and radially inward from the purge openings. A portion of a rotor blade is positioned radially outward from the purge openings. A mass flow rate of cooling air through the purge openings is selected based on a radial location of the purge openings to create an air barrier between a radially inner side of the purge openings and a radially outer side of the purge openings.

Abstract: A gas turbine engine has a fan drive turbine driving a gear reduction, the gear reduction, in turn, driving a fan rotor, the fan rotor delivering air into a bypass duct as bypass air and into a compressor section as core flow. A forward bearing is positioned between the gear reduction and the fan rotor and supports the gear reduction. A second bearing is positioned aft of the gear reduction and supports the gear reduction. The second bearing is a thrust bearing. A fan drive turbine drive shaft drives the gear reduction. The fan drive turbine drive shaft has a weakened link which is aft of the second bearing such that the fan drive turbine drive shaft will tend to fail at the weakened link, and at a location aft of the second bearing.

Abstract: An assembly is provided that includes a shaft, a bearing, a stator seal element, a rotor seal element and a shield. The shaft extends along an axis. The bearing supports the shaft and receives lubrication fluid. The stator seal element circumscribes the shaft. The rotor seal element is mounted on the shaft axially between the bearing and the stator seal element. The rotor seal element forms a seal with the stator seal element. The shield substantially prevents the lubrication fluid from traveling axially away from the bearing onto the rotor seal element.

Abstract: An airfoil for a gas turbine engine according to an example of the present disclosure includes, among other things, an airfoil section that has an internal wall and an external wall. The external wall defines pressure and suction sides that extends in a chordwise direction between a leading edge and a trailing edge, a first impingement cavity and a second impingement cavity bounded by the external wall at a leading edge region that defines the leading edge. A first crossover passage within the internal wall is connected to the first impingement. The first crossover passage defines a first passage axis that intersects a surface of the first impingement cavity. A second crossover passage within the internal wall is connected to the second impingement cavity. The second crossover passage defines a second passage axis that intersects a surface of the second impingement cavity.

Abstract: An apparatus for backward flow forming a material may comprise a mandrel having a headstock at a proximate end of the mandrel, the mandrel configured to rotate about an axis, a plurality of rollers disposed radially outward of the mandrel configured to exert force on the material to form a work piece at a plastic deformation zone, wherein the work piece flows from the plastic deformation zone between the plurality of rollers and the mandrel toward a distal end of the mandrel, and a catcher, coaxial to the mandrel, and removably coupled to the work piece at a traveling end of the work piece.

Abstract: A seal system has: a first member; a seal carried by the first member and having a seal face; and a second member rotatable relative to the first member about an axis. The second member has: a seat on a first piece of the second member, the seat having a seat face in sliding sealing engagement with the seal face; and a radially outwardly closed collection channel for collecting centrifuged oil; a second piece encircling and attached to the first piece and having a circumferential array of apertures; and cooperating with the first piece to define a plenum; and a flowpath from the collection channel passing radially outward axially spaced from the seat face to cool the seat face and passing axially away from the seat face in the plenum.

Abstract: Oil slinger systems include a seal runner comprising an annular radial member having a radius (R) and an outer axially extending member having an axial length (L), wherein a proximal surface of the outer axially extending member comprising a plurality of helical grooves. Methods of radial convective cooling include pumping a cooling liquid through the oil slinger system and convectively cooling the oil slinger.

Abstract: 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 MoS2 as the solid lubricant, although others are also disclosed.

Abstract: An airfoil includes an airfoil section that has an airfoil wall that define leading and trailing ends and first and second sides that join the leading and trailing ends. The first and second sides span in a longitudinal direction between first and second ends. The airfoil wall circumscribes an internal core cavity. The airfoil section extends form a platform. The platform defines a shelf that extends forward from the first end at the leading end of the airfoil section to a platform leading edge, aft from the first end at the trailing end of the airfoil section to a platform trailing edge, laterally from the first end at the first side of the airfoil section to a first platform side edge, and laterally from the first end at the second side of the airfoil section to a second platform side edge. The internal core cavity extends from the airfoil section into the shelf.