Raytheon Technologies Corporation

United Technologies Corporation

Rtx Corporation

James Glaspey

Jason Husband

Aaron John Pepin

Matthew A Turner

Andrew G Alarcon

Barry M Ford

David A Welch

Christopher M Quinn

Nicholas D Stilin

Renee J Jurek

Scott A Smith

Brian Green

Renee J Behlman

A rotor assembly for a gas turbine engine includes a rotatable hub that has a main body and an array of annular flanges that extend about an outer periphery of the main body to define an array of annular channels. An array of airfoils are circumferentially distributed about the outer periphery. Each one of the airfoils includes an airfoil section extending from a root section received in the annular channels. Retention pins extend through the root section of a respective one of the airfoils and through the array of annular flanges to mechanically attach each root section to the hub.

Rotor assembly for gas turbine engines

A method for forming a blade for a gas turbine engine may include forming a suction side sheath and a pressure side sheath, a first cavity and a second cavity established on opposed sides of a rib, forming a structural core configured for positioning in an interior section of the blade between the suction side sheath and the pressure side sheath, the structural core including a first core member, a second core member and a root interconnecting the first and second core members, assembling the suction side sheath and the pressure side sheath with the structural core such that the first core member is positioned in the first cavity and such that the second core member is positioned in the second cavity, and securing the suction side sheath to the pressure side sheath to form the blade.

Fan blade having closed metal sheath

An airfoil for a gas turbine engine according to an example of the present disclosure includes, among other things, an airfoil section extending from a root section. The airfoil section includes a sheath that receives a core. The core includes first and second ligaments received in respective internal channels defined by the sheath, and each one of the first and second ligaments includes at least one interface portion in the root section dimensioned to receive a retention pin.

Hybrid airfoil for gas turbine engines

A platform for a gas turbine engine according to an example of the present disclosure includes, among other things, a platform body that extends in a circumferential direction between first and second sidewalls to define a gas path surface, and opposed rows of flexible retention tabs that extend from the platform body and are dimensioned to wedge the platform between adjacent airfoils.

Rotor assembly with structural platforms 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, and the airfoil section defines a pressure side and a suction side separated in a thickness direction. The airfoil section includes a metallic sheath that receives a composite core, and the core includes first and second ligaments received in respective internal channels defined by the sheath such that the first and second ligaments are spaced apart along the root section with respect to the chordwise direction. The first and second ligaments define respective sets of bores, and the respective sets of bores are aligned to receive a common set of retention pins.

Pinned airfoil 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

A rotor assembly for a gas turbine engine includes, among other things, a rotatable hub that has a metallic main body that extends along a longitudinal axis, and that has an array of annular flanges that extend about an outer periphery of the main body to define an array of annular channels along the longitudinal axis. Each of the annular channels receives a composite reinforcement member that extends about the outer periphery of the hub.

A blade for use in a gas turbine engine is disclosed. In various embodiments, the blade includes a pressure side sheath and a suction side sheath secured to the pressure side sheath. The pressure side sheath and the suction side sheath are configured to form a continuous sheath that wraps around an interior section of the blade.

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. The airfoil section includes a metallic sheath that receives a composite core. The core includes first and second ligaments received in respective internal channels defined by the sheath such that the first and second ligaments are spaced apart along the root section with respect to the chordwise direction. Each one of the first and second ligaments includes at least one interface portion in the root section, and at least one interface portion of the first ligament and the at least one interface portion of the second ligament define respective sets of bores aligned to receive a common retention pin.

A fan of a gas turbine engine is provided. The fan having: a plurality of fan blades secured to a rotor, each of the plurality of fan blades having an airfoil secured to the rotor at one end and a tip portion that is secured to a shroud that circumscribes the plurality of fan blades.

Fan assembly of a gas turbine engine with a tip shroud

Disclosed is a fan blade for a gas turbine engine, the fan blade having: an airfoil having a leading edge, a trailing edge, a tip, and a frangible strip connected to the blade tip and extending outwardly therefrom, the frangible strip being less resistant to plastic deformation than the fan blade tip.

Fan blade tip with frangible strip

A spacer assembly for a rotor assembly of a gas turbine engine includes an endwall segment having a non-axisymmetric flowpath surface, a first depression and a second depression. A perimeter of the flowpath surface includes a forward edge, an aft edge, a suction side edge and a pressure side edge. The first depression is formed along the flowpath surface adjoining the suction side edge, and the second depression is formed along the flowpath surface adjoining the pressure side edge.

Contoured flowpath surface

In one embodiment, a hub for a fan of a gas turbine engine is provided. The hub having: a plurality of attachment features located on an outer circumferential surface of the hub, wherein at least some of the plurality attachment features extend radially away from the outer circumferential surface and are axially aligned with each other and at least some of the plurality of attachment features extending radially away from the outer circumferential surface and are off set from each other, and wherein the plurality of attachment features have an opening configured to receive a portion of a pin; and wherein at least some of the plurality of attachment features are located on a forward leading edge of the hub.

Fan hub attachment for leading and trailing edges of fan blades

Anti-rotation tabs for the platforms in the fan section of a gas turbine engine are provided. The anti-rotation tabs interface with the trailing edge of the spinner, thereby preventing the platform from rotating and twisting.

Growing community of inventors

Farmington, CT, United States of America

James Glaspey