An airfoil assembly for a gas turbine engine includes: an inner support structure configured to circumscribe a longitudinal axis of the gas turbine engine; an outer support structure configured to circumscribe the longitudinal axis of the gas turbine engine, the outer support structure circumscribing the inner support structure; and a stage including a plurality of airfoils extending from the inner support structure towards the outer support structure, the plurality of airfoils including: a first airfoil defining a first sweep angle, a first axial position, and a first lean angle; and a second airfoil defining a second sweep angle, a second axial position, and a second lean angle, wherein the second sweep angle is different than the first sweep angle, the second axial position is different than the first axial position, the second lean angle is different than the first lean angle, or a combination thereof.

A nacelle for a gas turbine engine having a longitudinal centre line includes an intake lip disposed at an upstream end of the nacelle. The intake lip includes a crown and a keel. The crown includes a crown leading edge and the keel includes a keel leading edge. The crown leading edge and the keel leading edge define a scarf line therebetween. The scarf line forms a scarf angle (θ.sub.scarf) relative to a reference line perpendicular to the longitudinal centre line. A fan casing is disposed downstream of the intake lip and includes a casing leading edge. The casing leading edge defines a droop line normal to the casing leading edge. The droop line forms a droop angle (θ.sub.droop) relative to the longitudinal centre line. A relationship between the droop angle (θ.sub.droop) and the scarf angle (θ.sub.scarf) is given by: θ.sub.droop=θ.sub.scarf/1.5±1 degree.

Various aspects of the disclosure are directed to providing structures that define a radial gap between an impeller and a pump casing element that facilitates minimizing the movement of fluid into the radial gap in a manner that lessens the impact, and consequent degradation, of the inner surface of the pump casing element by movement of abrasive particulates out of the radial gap, which is accomplished by providing a suction inlet arrangement of an impeller and pump casing element that are angled from the eye of the impeller to the outer periphery of the impeller in a direction away from the back shroud or drive side of the impeller and toward a first end of the pump casing in which fluid is introduced into the pump casing.

A coupling has: a first coupler rotatable about an axis and defining first connections distributed about the axis; a second coupler defining second connections distributed about the axis, the second connections offset from the first connections; and segments distributed about the axis and extending radially from the first connections to the second connections, a segment of the segments having a first end engaging a first connection of the first connections and a second end engaging a second connection of the second connections, the first end circumferentially offset from the second end, a face of the segment abutting against a face of the first coupler when the segment is inserted into the first connection in a first orientation such that a penetration depth of the segment into the first connection in the first orientation is less than the penetration depth in a second orientation opposite the first orientation.

The present invention thus proposes an inlet cone for an aircraft turbine engine, comprising a frustoconical body and a tip made from elastically deformable material fixed to an end of smaller diameter of said body, the tip comprising a top configured to be situated on an axis of rotation of the cone and a fastening base for attachment on said end of said body. Said base extends in a connecting plane P. Said connecting plane P is inclined relative to said axis of rotation. Said base has a generally circular or oval shape. According to the invention, said connecting plane P is inclined relative to a transverse plane T perpendicular to said axis of rotation.

A turbine vane including at least one inner cavity including a plurality of deflectors which are carried by an inner face of the lower surface wall and by an inner face of the upper surface wall, wherein each deflector extends mainly in a transverse direction from the inner face of the lower surface wall or from the inner face of the upper surface wall, in the direction of the other one of the lower surface wall or the upper surface wall, and wherein the length of each deflector in the transverse direction is greater than half the transverse distance between the inner face of the lower surface wall and the inner face of the upper surface wall, on either side of the deflector.

A turbine shroud assembly for use with a gas turbine engine includes a turbine outer case, a blade track segment, and a carrier. The turbine outer case is arranged circumferentially around an axis. The blade track segment includes an arcuate runner that extends circumferentially partway around the axis to define a gas path boundary of the turbine shroud assembly and an attachment feature that extends radially outward from the runner. The carrier is configured to couple the blade track segment to the turbine outer case.

A turbine blade includes an aerofoil including a leading edge, a trailing edge, a first sidewall, a second sidewall, and an internal cooling circuit disposed within the aerofoil and configured to direct a cooling fluid within the aerofoil. The turbine blade includes at least one first recessed portion formed on the first sidewall proximal to a tip of the turbine blade. The first recessed portion is disposed proximal to and spaced apart from the trailing edge of the aerofoil. The first recessed portion includes a base surface, a first surface, and a second surface. The first recessed portion further includes at least one slot extending from the first surface to the internal cooling circuit. The slot is configured to allow a flow of the cooling fluid from the internal cooling circuit to the first recessed portion.

A nacelle for a gas turbine engine having a longitudinal centre line. The nacelle includes an air intake disposed at an upstream end of the nacelle. The air intake includes, in flow series, an intake lip, a throat and a diffuser. The diffuser further includes a diffuser angle (θ.sub.diff), indicating a degree of divergence of the diffuser relative to the longitudinal centre line. The diffuser angle (θ.sub.diff) is from about 0 degrees to about 12 degrees.

A nacelle for a gas turbine engine having a longitudinal centre line. The nacelle includes an air intake disposed at an upstream end of the nacelle. The air intake includes, in flow series, an intake lip, a throat and a diffuser. The nacelle further includes a protrusion extending radially inward from the air intake downstream of the intake lip. The protrusion extends circumferentially by a protrusion angle (θ.sub.p) with respect to the longitudinal centre line of the gas turbine engine.