An integrally bladed rotor includes a rotor disk, rotor blade, internal channel and air outlets. The rotor blade includes a body defined by a pressure surface and suction surface between a leading edge and a trailing edge. The internal channel extends radially from the rotor disk internally within the body of the rotor blade. The air outlets are formed proximate the leading edge and extend within the body of the rotor blade to the internal channel.

An integrally bladed rotor includes a rotor disk, rotor blade, internal channel and air outlets. The rotor blade includes a body defined by a pressure surface and suction surface between a leading edge and a trailing edge. The internal channel extends radially from the rotor disk internally within the body of the rotor blade. The air outlets are formed proximate the leading edge and extend within the body of the rotor blade to the internal channel.

A system is provided for a gas turbine engine. This engine system includes a flowpath wall, a takeoff conduit and a porous cover. The flowpath wall forms a peripheral boundary of an internal engine flowpath. The flowpath wall includes a takeoff port. The takeoff conduit includes an internal conduit passage fluidly coupled with the internal engine flowpath through the takeoff port. The takeoff conduit projects out from the flowpath wall. The porous cover for the internal conduit passage is disposed at the takeoff port.

A stator blade for a centrifugal compressor having a front portion configured to generate one or more strear-wise vortices in the gas flow around the stator blade in order to avoid and/or delay a detachment of the gas flow from the suction surface of the stator blade especially when the centrifugal compressor is not operating at its operational design speed.

An airfoil assembly for a turbine engine comprising an outer band, an inner band radially spaced inwardly from the outer band to define an annular region, and multiple airfoils circumferentially spaced within the annular region. Each corresponding airfoil of the multiple airfoils can project from a surface at a root and can further include an outer wall defining a pressure side and a suction side. A projection can extend upwardly from the surface on the pressure side and a valley can extend into the surface on the suction side to define a contour in the surface.

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.

The invention concerns a support casing (24) for a rotor (12) of a turbine engine such as a ducted fan turbojet engine used for propulsion of an aircraft. The casing (24) comprises: an outer annular wall (38) with an inner annular surface (44); an inner hub (40) able to support the rotor (12) of the axial turbine engine and comprising an outer annular surface (42); an annular passage (46) between the annular wall (38) and the inner hub (40); an annular row of arms (48) passing radially through the annular passage (46). Each arm (48) of the casing (24) comprises an orifice (50) arranged in the annular passage (46) radially at the level of one of said annular surfaces (42; 44). Inserts are fitted to the orifices (50) to control the flow passing through.

A strut for a gas turbine engine includes an airfoil section extending in a spanwise direction between a first platform and a second platform, extending in a chordwise direction between a leading edge and trailing edge to define a chord length, and extending in a thickness direction between a first side and a second side to define a chord width. Exterior surfaces of the airfoil section define a leading portion between the leading edge and a widest location of the airfoil section relative to the thickness direction, and a trailing portion between the widest location and the trailing edge. The exterior surfaces establish a respective exterior contour for each span position between a 0% span position and a 100% span position. The exterior surfaces define a plurality of dimples in the leading portion.

Air inlet duct of a turbine engine, in particular an aircraft turbine engine comprising a gas generator, which extends axially between the air inlet and the gas generator and has a first axial wall part and a second wall part which is angularly offset with respect to the first part, which duct is capable of causing, in a shedding region, shedding of the boundary layer formed by an air flow along the wall of the duct; and a device for controlling said shedding of the boundary layer, characterised in that the control device comprises an air-blowing pipe which opens via at least one air-injection opening which is directly upstream of the shedding region, the blowing pipe being connected to an air intake positioned upstream of said air-injection opening or in the shedding region and comprising an air compressor means between the air intake and the air-injection opening.

A turbine blade according to the invention includes a first wall surface facing a cooling passage through which cooling air flows; a second wall surface facing a working fluid passage through which a working fluid flows; a communication hole establishing communication between the cooling passage and the working fluid passage; and a projection provided on the downstream side of the flowing direction of the cooling air in the opening of the communication hole formed in the first wall surface, the projection protruding from the first wall surface toward the cooling passage. The projection includes a slope section formed on the first wall surface that ascends in the direction opposite the flowing direction of the cooling air toward the cooling passage and a curved section formed as a convex section on the side of the cooling passage, the curved section extending along the opening in the form of an arc.