A system is provided for an aircraft propulsion system. This system includes an inlet lip, a bulkhead and a piccolo tube for a thermal anti-icing system. The inlet lip extends circumferentially about an axial centerline. The bulkhead extends circumferentially about the axial centerline. The bulkhead is configured with the inlet lip to form a cavity axially between the inlet lip and the bulkhead. The piccolo tube extends circumferentially about the axial centerline within the cavity. The piccolo tube is configured with an elliptical cross-sectional geometry.

A liquid rocket engine integrates tap-off openings at a combustion chamber wall to direct exhaust from the combustion chamber to a tap-off manifold that provides the exhaust to one or more auxiliary systems, such as a turbopump that pumps oxygen and/or fuel into the combustion chamber. The tap-off opening passes through a fuel channel formed in that combustion chamber exterior wall and receives fuel through a fuel opening that interfaces the fuel channel and tap-off opening. The tap-off manifold nests within a fuel manifold for thermal management. The fuel channel directs fuel into the combustion chamber through fuel port openings formed in the combustion chamber, the fuel port openings located closer to a headend of the combustion chamber than the tap-off openings.

An injector for a combustor of a gas turbine engine is provided with a plurality of first vanes radially arrayed about a central axis of the injector and a plurality of second vanes radially arrayed about the central axis of the injector and disposed radially inward of the plurality of first vanes. A plurality of fuel injection holes are disposed nearer to a trailing edge than to a leading edge of the second vanes for injecting fuel into compressed air passing through over the second vanes. The trailing edge of each of the second vanes includes a non-planar profile configured to induce turbulence in the compressed air to thereby mix the fuel with the compressed air and reduce the surfaces on which undesirable flame anchoring may occur.

An acoustic liner for a turbine engine, the acoustic liner includes a support layer that includes a set of partitioned cavities with open faces, a facing sheet operably coupled to the support layer such that the facing sheet overlies and closes the open faces, a set of perforations located in the facing sheet and in fluid communication with a corresponding one of the cavities to form a set of acoustic resonators, and at least a subset of the perforations have an axially-oriented, relative to the axial flow path, inlet.

An improved system, apparatus and method may comprise an inner liner and an outer liner extending circumferentially around an engine centerline axis thereby forming a combustion chamber therebetween. The system, apparatus, and method may include a front end extending between the inner and outer liners and having a plurality of fuel nozzle ports configured to each receive a fuel nozzle. First and second outer plunged inlet holes may be formed in the outer liner and extending in a first radial direction toward the engine centerline axis and into at least a portion of the combustion chamber. First and second inner plunged inlet holes may be formed in the inner liner and extending in a second radial direction away from the engine centerline axis and into at least a portion of the combustion chamber.

Disclosed herein is a gas turbine blade. The gas turbine blade includes a turbine blade (33) provided in a turbine, and film cooling elements (100), each including a cooling channel (110) for cooling of the turbine blade (33), an outlet (120) through which cooling air is discharged, and a plurality of ribs (130), wherein the outlet (120) extends from a longitudinally extended end of the cooling channel (110) to an outer surface of the turbine blade (33) and has a width increased from one end of the cooling channel (110) to the outer surface of the turbine blade (33), and the ribs (130) face each other on inner walls of the outlet (120).

A core element of an investment core for use in a casting process used to produce an airfoil includes an investment core body, an extension connected to and protruding from the investment core body, and a connection portion connected to the investment core body and to the extension. The investment core body comprises a ceramic material. A shape of the extension comprises a tube with a centerline axis passing through a center of the extension. A shape of a cross-section of the extension taken along a plane perpendicular to the extension centerline axis comprises an ellipse. The extension is connected to the investment core body by the connection portion.

Disclosed herein is a turbine blade having a leading edge, a trailing edge, a pressure side, and a suction side and having a cooling passage formed therein. The turbine blade includes a plurality of cooling holes formed on the pressure side or the suction side. Each of the cooling holes includes, a through-hole having a circular cross-section and angled outwardly toward the pressure side or the suction side from the cooling passage, a circular sink formed concavely with respect to the pressure side or the suction side in the vicinity of an upstream side of an exit of the through-hole, and an elliptical sink formed concavely with respect to the pressure side or the suction side in the vicinity of a downstream side of the exit of the through-hole.

A gas turbine engine includes an inlet duct that is formed with a generally elliptical shape. The inlet duct includes a vertical centerline and a fan section that has an axis of rotation. The axis of rotation is spaced from the vertical centerline and is disposed within an inlet duct orifice.

An airfoil includes: an airfoil body extending from a radially inner root portion to a radially outer tip portion, extending from a pressure side to a suction side, and from a leading edge to a trailing edge portion. The trailing edge portion includes: a linear pressure side portion; a linear suction side portion; a trailing edge tip portion; and at least one elliptical portion disposed between the trailing edge tip portion and at least one of the linear pressure side portion and the linear suction side portion. The trailing edge portion may also include a circular trailing edge tip portion.