Disclosed is a turbomachine including a stator, a rotor rotatable about an axis of rotation, and an annulus for carrying a core flow, the annulus having a side wall on the stator and a side wall on the rotor, at least one airfoil array having a plurality of airfoils being disposed in the annulus. In a departure from an ideal aerodynamic annulus contour, a radial annulus enlargement begins upstream of the airfoils and extends downstream up to an aft portion of the airfoil array that follows the ideal aerodynamic annulus contour. Also disclosed is an airfoil for such a turbomachine.

A windage shield (200, 200) for mounting on a fan disc (100, 100) of a fan (23) of a gas turbine engine (10), the windage shield (200, 200) comprising: a fan disc contacting portion (215, 215) adapted to contact and structurally support a rear portion of the fan disc (100, 100); wherein the fan disc contacting portion (215, 215) includes one or more stiffening elements (205, 205) to locally increase the hoop stiffness of the windage shield (200, 200).

A fastener assembly for a gas turbine engine, and method of assembly, includes a first body having a first surface and a recess communicating through the first surface. The recess may be defined by a bottom surface and a side face spanning between the first and bottom surfaces. A shank of the assembly is generally engaged between the first body and a second body and includes opposite first and second end portions. The first end portion is located in the recess and the second end portion is engaged to the second body. A filler of the assembly is generally located in the recess to cover the first end portion. To reduce windage, the filler has an outer surface that is substantially flush with the first surface.

A nosecone assembly having an axially extending centerline is provided. The assembly includes a nosecone body and at least one access panel. The nosecone body has at least one wall that defines an interior cavity. The wall has an interior surface contiguous with the interior cavity, and at least one window aperture extending through the wall. The access panel has first and second face surfaces. The access panel is attached to the wall interior surface within an attachment region that includes first and second attachment region portions partially contiguous with one another. The first and second attachment region portions define an interior unattached region, and the interior unattached region is aligned with the window aperture.

A fan assembly for use in a gas turbine engine of an aircraft includes a fan disk having a number of fan blades and a windage shield coupled to the fan disk to move therewith. The fan assembly supplies air for use in the engine. The windage shield rotates with the fan disk during operation of the gas turbine engine and directs air supplied by the fan blade.

A turbine engine having a rotor rotatable about a rotational axis, a stator, a plurality of circumferentially spaced bleed air passages, and an inducer. The plurality of circumferentially spaced bleed air passages being located between an axially adjacent set of vanes of the stator and blades of the rotor. The inducer including a nozzle passage fluidly coupling a nozzle inlet of the inducer to a nozzle outlet of the inducer.

A gas turbine engine includes a rotating stage of blades. A circumferential array of blade outer air seal segments are arranged radially outward of the blades. Adjacent blade outer air seal segments provide a circumferential gap. Facing ends of the adjacent blade outer air seal segments include surfaces. A gap seal engages the surfaces and obstructs the circumferential gap. A biasing member is configured to urge the gap seal radially inward toward the surfaces.

A method for installing a probe assembly in a case of a gas turbine engine the method including installing a first portion of the probe assembly within a first section of the case, and installing a second portion of the probe assembly within a second section of the case. A case assembly within a gas turbine engine the case assembly including a case in at least one of a compressor and a turbine, and a probe assembly. The probe assembly including a first portion positioned within a bore of the case, and a second portion positioned within an inset of the case, the bore having a smaller diameter than the inset.

A gas turbine engine according to the present disclosure includes a first component, a second component coupled to the first component, and a pilot unit. The pilot unit provides means for maintaining a pilot-setting force between the first and second component to retain alignment of the first component with the second component.

A gas turbine engine includes a blade having a tip, a blade outer air seal operatively connected to a case assembly, and an active clearance control system disposed on the case assembly. The active control system includes an actuator assembly. The actuator assembly includes a motor assembly and a shaft. The shaft has a shaft body that extends between a first end that is operatively connected to the motor assembly and a second end that is operatively connected to the blade outer air seal.