A gas turbine engine component has a component body configured to be positioned within a flow path of a gas turbine engine having an external pressure, and wherein the component body includes at least one internal cavity having an internal pressure. At least one inlet opening is formed in an outer surface of the component body to direct hot exhaust gas flow into the at least one internal cavity, and there is at least one outlet from the internal cavity. The internal pressure is less than an inlet external pressure at the inlet opening and the internal pressure is greater than an outlet external pressure at the outlet opening to controllably ingest hot exhaust gas via the inlet opening and expel the hot exhaust gas via the outlet opening to maintain a laminar boundary layer along the outer surface of the component body.

A gas turbine engine component has a component body configured to be positioned within a flow path of a gas turbine engine having an external pressure, and wherein the component body includes at least one internal cavity having an internal pressure. At least one inlet opening is formed in an outer surface of the component body to direct hot exhaust gas flow into the at least one internal cavity, and there is at least one outlet from the internal cavity. The internal pressure is less than an inlet external pressure at the inlet opening and the internal pressure is greater than an outlet external pressure at the outlet opening to controllably ingest hot exhaust gas via the inlet opening and expel the hot exhaust gas via the outlet opening to maintain a laminar boundary layer along the outer surface of the component body.

Technologies are described herein for a drag recovery scheme using a boundary layer bypass duct system. In some examples, boundary layer air is routed around the intake of one or more of the engines and reintroduced aft of the engine fan in the nozzle duct in a mixer-ejector scheme. Mixer-ejectors mix the boundary layer flow to increase mass flow.

A steam turbine having an inflow ring channel which is connected to an inflow connecting piece in terms of flow technology. The inflow connecting piece is designed in such a way that an incoming flow is first slowed down, subsequently accelerated and simultaneously deflected.

An apparatus and method of minimizing airfoil boundary layer separation utilizing at least one bleed inlet disposed on the outer wall of the airfoil, such as the suction side, having at least one channel disposed within an interior of the airfoil providing fluid communication between the bleed inlet and the tip of the airfoil. Bleed gas drawn through the bleed inlet and provided to the tip can pressurize a seal disposed at the tip. Additionally, a flow control device can be disposed within the channel to control or meter the rate at which gas is bled into the channel.

An aircraft inlet comprising an annular air intake structure defined by an inner surface and an outer surface, a plurality of perforations formed in the outer surface of the aircraft inlet, a plenum chamber situated within the aircraft inlet and configured to receive air entering the plurality of perforations, and/or a plurality of sensors disposed about the outer surface of the aircraft inlet, each sensor associated with a region of the plurality of perforations. Each sensor may comprise a hot film anemometer. The aircraft inlet may further comprise a regulator coupled at one end to the plenum chamber and at another end to a pump, wherein the regulator regulates a suction produced by the pump.

A vane for use in a gas turbine engine includes an airfoil having a suction side and a pressure side, and a transpirational flow acoustic liner disposed in the airfoil. The liner includes a face sheet defining a portion of an outer surface of the airfoil and having a plurality of first apertures, a segmented member coupled to the face sheet, and a backing sheet coupled to the segmented member the segmented member such that the segmented member is positioned between the face sheet and the backing sheet. The segmented member includes a plurality of chambers in fluid communication with the outer surface via the plurality of first apertures. The backing sheet has a plurality of second apertures.

A centrifugal compressor diffuser (42) includes a plurality of diffuser flow passages (22) extending through an annular diffuser housing (20) and circumferentially bounded by diffuser vanes (23) and axially bounded by forward and aft walls (101, 100). A diffuser boundary layer bleed (96) for the passages may include boundary layer bleed apertures (106) or slots (130) disposed through the forward wall (101) and a downstream facing wall (142) canted at an acute cant angle to a downstream diffuser airflow direction (103) in the passages. Diffuser bleed flow (112) is bled from a diffuser boundary layer. Boundary layer bleed apertures can be located downstream of throat sections (28) of the flow passages near pressure sides of the vanes. A centrifugal compressor (18) may include the diffuser surrounding an annular centrifugal compressor impeller (32) and apparatus for flowing impeller bleed flow (102) from a radial clearance between an impeller tip (36) and a diffuser annular inlet (27) with diffuser bleed flow either mixed or separately to cool a turbine (16).

A component stage for a turbomachine includes a component section. The component section includes a flowpath surface at least partially exposed to a core air flowpath defined by the turbomachine, when the component stage is installed in the turbomachine. The component further includes a plurality of sequentially arranged riblets on the flowpath surface, the plurality of sequentially arranged riblets customized for an anticipated location of the flowpath surface within the turbomachine by defining one or both of a non-uniform geometry or a non-uniform spacing.

A compressor airfoil in a gas turbine engine is presented. Opposed pressure and suction sides are joined together at chordally opposite leading and trailing edges. The pressure and suction sides extend in a span direction from a root to a tip of the airfoil. A leading edge sweep angle is defined relative to a tangent to the airfoil and flow velocity vector at a point on the leading edge. A leading edge dihedral angle is defined relative to the tangent to the airfoil and a vertical at the point on the leading edge. A ratio of the leading edge sweep angle to the leading edge dihedral angle being smaller than 1. A method of forming such airfoil is also presented.