Abstract: A brush seal assembly is adapted to be operatively mounted between a stator and a rotor of a gas turbine engine. The brush seal assembly comprises a housing and a seal operatively coupled to the housing. The seal comprises at least two annular brushes each having a plurality of bristles extending away from the housing. The at least two annular brushes and the housing defines a cavity therebetween. The housing defines at least one circumferential surface within the cavity. The housing comprises at least one discontinuity defined in the at least one circumferential surface. The at least one discontinuity disrupts a continuity of annularity of the circumferential surface.

Abstract: A gas turbine engine includes a low pressure compressor and a low pressure turbine, with a low pressure shaft coupled to the low pressure turbine through a turbine extension and coupled to the low pressure compressor through a compressor extension. A forward bearing assembly supports the low pressure shaft within a compressor section and at a location aft of the compressor extension and an aft LP bearing assembly supports the low pressure shaft within a turbine section at a location aft of the turbine extension.

Abstract: A blade includes an airfoil, a stationary portion coupled to a radially inner end of the airfoil, and a leakage flow guide vane assembly coupled to the stationary portion. The leakage flow guide vane assembly includes a plurality of passages defined therein. The passages are oriented to induce a swirl velocity to a working fluid flowing through the passages.

Abstract: A hydrostatic seal assembly includes a seal support, a seal shoe integral to the seal support and radially movable relative to the seal support. A plurality of seal teeth are located at a radially inboard surface of the seal shoe. The seal shoe is formed from a first material and the radially inboard surface is formed from a second material softer than the first material such that the radially inboard surface is abradable in the event of a rub between the radially inboard surface and an adjacent rotating component.

Abstract: The present disclosure is directed to a gas turbine engine including a turbine section including a first rotating component interdigitated along a longitudinal direction with a second rotating component. The first rotating component and the second rotating component are each coupled to a speed reduction assembly in counter-rotating arrangement. The first rotating component comprising an outer shroud and a plurality of outer shroud airfoils extended inward along a radial direction from the outer shroud. A connecting member couples the outer shroud to a radially extended first rotor. The second rotating component comprising an inner shroud and a plurality of inner shroud airfoils extended outward along the radial direction from the inner shroud, the plurality of inner shroud airfoils in alternating arrangement along the longitudinal direction with the plurality of outer shroud airfoils.

Abstract: A housing for an air cycle machine includes a first housing portion adjacent a first end of the housing, and a second housing portion adjacent a second end of the housing and attached to the first housing portion. A first duct in the first housing portion is adjacent the first end of the housing with an air inlet and an air outlet. A second duct in the second housing portion is adjacent the second end of the housing with an air inlet and an air outlet. The housing further includes a bolt hole positioned in the second end of the housing, and a rib extending between the first housing portion and the second housing portion. The rib is aligned with the bolt hole.

Abstract: Disclosed herein is a disk assembly for a gas turbine compressor, which comprises a partition wall formed to partition a space between disks for a gas turbine compressor to optimize a cooling fluid path.

Abstract: An airport having an exterior wall including a plurality of spaced layers for improved cooling and lifetime is disclosed. The airfoil and exterior wall are made by additive manufacturing. The exterior wall includes an exterior layer, an intermediate layer, and an interior layer each separated from adjacent layers by a plurality of standoff members; a plurality of first cooling chambers between the exterior and intermediate layers, the chambers partitioned by a first partitioning wall; a plurality of second cooling chambers between in the intermediate and interior layers, the chambers partitioned by a second partitioning wall; a thermal barrier coating on the exterior layer; a plurality of impingement openings in the intermediate layer and a second plurality of impingement openings in the interior layer; and a plurality of cooling passages in the exterior layer. The exterior layer may also include plateaus on an exterior face through which the cooling passages extend.

Abstract: A turbine vane made from ceramic matrix composite material and adapted for use in a gas turbine engine. The turbine vane is made from ceramic matrix composite materials with three-dimensionally woven or braided reinforcing tows and includes an inner band, an outer band, and an airfoil that extends between the inner band and the outer band.

Abstract: In an axial flow fan according to the present invention, a plurality of blades rotate about a rotation axis of the blades to convey a fluid. In the axial flow fan, the plurality of blades each have a leading edge at a leading side in a rotational direction, a trailing edge at a trailing side in the rotational direction, and an outer peripheral edge connecting the leading edge and the trailing edge. The leading edge of one of the plurality of blades and the trailing edge of another blade adjacent to the leading edge of the blade in the rotational direction are connected by a plate-shaped connection portion. The plurality of blades each have at least one plate-shaped reinforcement rib extending from a periphery of the rotation axis toward the outer peripheral edge of the blade.

Abstract: A turbomachine including a housing having an inlet end opposite and outlet end along a longitudinal axis of the housing, a shaft assembly provided within the housing, the shaft assembly extending from the inlet end to the outlet end, a rotor having at least one rotating impeller extending radially outward from the shaft assembly, and a return channel vane hub extending radially outward from the shaft assembly, the return channel vane hub includes at least one return channel vane extend therefrom, the at least one return channel vane comprising a body having a leading edge and a trailing edge, the leading edge is twisted and extended past an outer edge of the return channel vane hub, and the trailing edge is bowed outwardly.

Abstract: The invention relates to a flow rectifier for an axial fan as well as to an axial fan having a flow rectifier. A flow rectifier according to the invention for an axial fan with a hub and at least one vane is characterized in that the at least one vane can be operatively connected to the hub so as to be detachable. Since the vane and the hub can be connected so as to be detachable, the flow rectifier has a modular structure.

Abstract: A component for a gas turbine engine includes at least one airfoil that has a radially inner end and a radially outer end. A platform includes a radially outer surface that is attached to the radially inner end of the airfoil. A radially inner side of the platform includes a forward surface and an aft surface. A pocket is recessed into at least one of the forward surface and the aft surface. A cover plate covers the pocket and a seal is attached to the cover plate.

Abstract: A system, method, and apparatus for tuning an operating frequency of turbines blades may be provided. Multiple blades having tip shrouds may be provided. The blades may extend radially outward from an axis of rotation of the blades. A tip shroud of each blade may be located at a radial end of the blades. The tip shrouds may have side faces that interlock with faces of adjacent tip shrouds. In addition, some tip shrouds may have one or more gaps between some side faces of consecutive tip shrouds.

Abstract: A device for testing a rotor blade of an aircraft. The device is configured to receive and secure a shaft of the rotor blade. Once secured, the device is configured to allow for inputting combined loads and bending moments into the rotor blade to simulate the rotor blade behavior in flight. A testing assembly is also disclosed that includes the device and the rotor blade and methods of use that provide for testing the rotor blade.

Abstract: A component retaining assembly includes a housing including a circumferential slot. A component is mated to the housing. The component includes a first face including at least two tabs. The at least two tabs extending at outward from the first face at least partially past a portion of the circumferential slot. A retaining ring is disposed within the circumferential slot and abuts the first face of the component. The at least two tabs overlap a portion of the retaining ring. A gas turbine engine and a method are also disclosed.

Abstract: An apparatus and system for a marine propeller assembly are provided. An aft retention member that may be used with the marine propeller assembly includes a base, an opposing nose and a conic body extending therebetween along a centerline normal to the base. The aft retention member also includes at least one protuberance extending radially away from a surface of the conic body. The protuberance extends axially along a surface of the aft retention member from the base arcuately convergent to a predetermined point between the base and a tip of the nose.

Abstract: A thermal barrier attachment seal for a gas turbine engine component includes a spine having a first surface and a second surface facing opposite the first surface. A first layer of insulation is provided on the first surface. A second layer of insulation is provided on the second surface to provide a thermal barrier seal between a first gas turbine engine component and a second gas turbine engine component. A fan section for a gas turbine engine is also disclosed.

Abstract: A turbine rotor blade includes an airfoil portion having an airfoil defined by a pressure surface and a suction surface; and a squealer rib on a tip surface of the turbine rotor blade, the squealer rib extending from a leading-edge side (toward a trailing-edge side. The squealer rib has a ridge extending in an extending direction of the squealer rib. The turbine rotor blade is configured to provide a clearance between the tip surface of the turbine rotor blade and an inner wall surface of a casing of a turbine such that the inner wall surface of the casing of the turbine faces the tip surface of the turbine rotor blade and the clearance has a local minimum value on the ridge. The clearance is greater than the local minimum value at both sides of the ridge in a width direction of the squealer rib.

Abstract: An aerofoil member that extends between radially inner and radially outer endwalls which define a gas annulus of the compressor. Aerofoil member has a leading edge, a trailing edge, a pressure surface and a suction surface such that successive cross-sections through the aerofoil member transverse to the radial direction provide respective aerofoil sections. The external shape of the aerofoil member is defined by the stacking of the aerofoil sections on a stacking axis which passes through a reference point common to each aerofoil section. The aerofoil member has lean produced by the projection of the stacking axis onto a plane normal to the engine axis intersecting a first one of the endwalls at an angle of from 5° to 25° to the circumferential direction direction such that the pressure surface faces the first endwall.