Raytheon Technologies Corporation

United Technologies Corporation

Rtx Corporation

David A Welch

Jason Husband

Ioannis Alvanos

Thomas J Robertson, Jr

Mark W Costa

Frederick M Schwarz

James Glaspey

Phillip Alexander

Andre Herman Troughton

Michael Raymond LaFavor

William Meyst

Jesse Ian Berube

Michael Raymond Lafavor

Michael W Trudnak

Brian J Breslin

Christopher Durkee

Jr Thomas J Robertson

A fan blade airfoil and nose cone with leading edge fillet for debris deflection including a fillet formed in the nose cone and inserted into a pocket formed in the fan blade, the filet configured to create a deflection air stream; the deflection air stream being configured to manipulate a fan inlet air flow such that any debris entrained in a fan inlet air flow has a trajectory line directed away from a core flow of a gas turbine engine and toward a bypass flow of the gas turbine engine.

Fan blade assembly with leading edge fillet

A fan exit guide vane with a load member and cooling cavity including an inner attachment region opposite an outer attachment region; a load member cavity formed within the fan exit guide vane extending spanwise through the fan exit guide vane from the inner attachment region to the outer attachment region; the load member extending through the load member cavity beyond each of the inner attachment region and the outer attachment region of the fan exit guide vane; and a cooling cavity formed within the fan exit guide vane extending spanwise through the fan exit guide vane between the inner attachment region and the outer attachment region.

Fan load tension structural guide vane with bypass cooling cavities

A blade fabrication method is provided and includes additively manufacturing a core, securing the core to a mandrel, electroforming a leading edge sheath directly onto the core and the mandrel and removing the mandrel from the core and the leading edge sheath.

Composite fan blade

An airfoil assembly for a gas turbine engine according to an example of the present disclosure includes, among other things, an airfoil that has an airfoil section extending from a root section. The airfoil section extends between a leading edge and a trailing edge in a chordwise direction and extends between a tip portion and the root section in a radial direction, and the airfoil section defining a pressure side and a suction side separated in a circumferential direction. A platform includes a first and second platform portions attached to the root section. Each of the first and second platform portions includes a shoe section and a platform section extending in the circumferential direction from the shoe section to establish a gas path surface, and the shoe sections of the first and second platforms are circumferentially arranged on opposed sides of the root section to capture the root section in a root cavity established between the shoe sections. A method of assembly is also disclosed.

Airfoil with integral platform for gas turbine engines

An airfoil for a gas turbine engine according to an example of the present disclosure includes, among other things, an airfoil section that extends from a root section. The airfoil section extends between a leading edge and a trailing edge in a chordwise direction and extends between a tip portion and the root section in a radial direction. The airfoil section defines a pressure side and a suction side separated in a thickness direction, and the airfoil section includes a metallic sheath that defines an internal cavity receiving a composite core. The root section defines at least one bore dimensioned to receive a retention pin. At least one damping element is received in the internal cavity and that selectively causes the airfoil section to stiffen.

Rotor assembly with active damping for gas turbine engines

An airfoil includes a leading edge, a trailing edge, a pressure side extending between the leading and the trailing edge, and a suction side that is disposed opposite the pressure side extending between the leading edge and the trailing edge. The airfoil further includes a spar and a structural member. The spar defines the leading edge. The structural member is bonded to the spar.

Fan blade with structural spar and integrated leading edge

An airfoil assembly for a gas turbine engine according to an example of the present disclosure includes, among other things, an airfoil including an airfoil section extending from a root section. The airfoil section extends between a leading edge and a trailing edge in a chordwise direction and extending between a tip portion and the root section in a radial direction. The airfoil section defines a pressure side and a suction side separated in a circumferential direction. A root mount includes first and second mount members secured to circumferentially opposed sides of the root section. Arcuate circumferential faces of the first and second mount members are dimensioned to pivotably mount the root section to a hub in an installed position. A method of assembly is also disclosed.

Airfoil attachment for gas turbine engines

A gas turbine engine comprises a fan rotor having fan blades received within an outer nacelle, and the outer nacelle having an inner surface. At least a portion of the nacelle inner surface extends radially inwardly to be radially inward of an outer diameter of the fan blades. The inner surface of the nacelle is formed with a trench, which extends into the inner surface to a radially outer extent that is spaced radially outward of the outer diameter of the fan blades.

Gas turbine engine with short inlet and blade removal feature

A fan blade includes a lattice structure, a fill, and a structural wrap. The lattice structure defines an airfoil body that has a cavity. The lattice structure defines at least one spar that extends radially from a root of the airfoil body towards a tip of the airfoil body. The fill is disposed within the cavity. The structural wrap is disposed about the lattice structure. The structural wrap defines a plurality of plies that extend from a leading edge of the airfoil body towards a trailing edge of the airfoil body.

Micro lattice hybrid composite fan blade

A cartridge for a fan case of a gas turbine engine includes an inlet acoustic liner section integrated with a thermally conforming liner section.

Thermally conforming acoustic liner cartridge for a gas turbine engine

A nacelle for a gas turbine engine, the nacelle includes an inlet axially forward of a fan section. The inlet includes an interior surface. A leading edge includes an inlet cavity. A first conduit communicates airflow to the inlet cavity for preventing accretion of ice on the leading edge. An inlet duct is open to the interior surface aft of the inlet cavity. A second conduit is in fluid communication with the first conduit and the inlet duct. The second conduit communicates airflow to the inlet duct for generating film heating airflow over the interior surface for preventing ice accretion. A gas turbine engine is also disclosed.

Short inlet with integrated liner anti-icing

A heat exchanger manifold for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a manifold body extending between a first face and a second face, a first seal land defining a first fluid port extending through the manifold body between the first face and the second face, and a first seal received within the first fluid port.

Gas turbine engine heat exchanger manifold

A gas turbine engine includes a fan forward of a primary flowpath inlet. The fan includes multiple fan blades distributed radially about, and connected to, a hub. A fan nacelle is positioned radially outward of the fan and includes an inner diameter. The inner diameter is sloped relative to an engine axis such that a forward portion of the fan nacelle has a smaller inner diameter than an aft portion of the fan nacelle. A fan blade spacer is disposed between a radially inward facing surface of at least one fan blade and a radially outward facing surface of the hub. The fan blade spacer has a radial thickness at least equal to a slope drop of the fan nacelle.

Fan blade removal feature for a gas turbine engine

A gas turbine engine includes a liner disposed around a flowpath. The liner has a forward end, a radially outer surface, and a radially inner surface. A hole extends axially into the forward end of the liner between the radially outer surface and the radially inner surface, and an engagement member is partially disposed in the hole and extends axially forward from the forward end of the liner.

Growing community of inventors

Quaker Hill, CT, United States of America

David A Welch