Abstract: A gas turbine engine including: a turbomachine having a compressor section, a combustion section, and a turbine section arranged in serial flow order; a fan defining a fan axis and comprising a plurality of fan blades rotatable about the fan axis; and a pitch change mechanism operable with the plurality of fan blades, the pitch change mechanism including a plurality of linkages, the plurality of linkages including a first linkage coupled to a first fan blade of the plurality of fan blades and a second linkage coupled to a second fan blade of the plurality of fan blades; and a non-uniform blade actuator system operable with one or more of the plurality of linkages to control a pitch of the first fan blade relative to a pitch of the second fan blade.

Abstract: A gas turbine engine includes a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order. A fan defining a fan axis and comprising a plurality of fan blades is rotatable about the fan axis. A pitch change mechanism is operable with the plurality of fan blades to control a pitch of the plurality of fan blades. A counterweight is connected to each respective fan blade. A spring mechanism is operably associated with each fan blade of the plurality of fan blades to create a spring twisting moment on the respective fan blade that is counter to a centrifugal twisting moment of the respective fan blade.

Abstract: A rotor assembly is provided, along with gas turbine engines for its use. The rotor assembly may include a spool defining a plurality of apertures arranged in a first row and spaced circumferentially around the spool, wherein each aperture of the plurality of apertures extends through the spool from a radially inward-facing surface to a radially outward-facing surface; and a blade assembly comprising at least two blades connected to each other via a platform, wherein each blade extends through a respective aperture.

Abstract: A composite airfoil having a non-metallic leading edge protective wrap is provided. In one aspect, the airfoil has a composite core having a pressure sidewall and a suction sidewall each extending between a core leading edge and a core trailing edge. A leading edge protective wrap protects the core leading edge and includes a trailing wrap and a leading wrap. The trailing wrap wraps around the core leading edge and is connected to the composite core. The leading wrap wraps around the core leading edge and is connected to the trailing wrap. The trailing and leading wraps both have leading edges that are spaced from one another. A filler is positioned between the leading edges of the trailing and leading wraps. A protective nose is connected to the leading edge of the leading wrap. The components of the leading edge protective wrap are formed of non-metallic materials.

Abstract: Hybrid composite components, such as gas turbine engine containment assemblies, a hybrid composite component having an annular composite shell and an annular metallic shell joined with the composite shell. A containment assembly including a containment case extending along an axial direction about a longitudinal centerline of the gas turbine engine. The containment case has an inner surface and an outer surface spaced apart along a radial direction and includes a first composite shell joined with a metallic shell. The metallic shell defining a first portion of the inner surface and the first composite shell defining a second portion of the inner surface.

Abstract: A gas turbine engine including: a turbomachine having a compressor section, a combustion section, and a turbine section arranged in serial flow order; a fan defining a fan axis and comprising a plurality of fan blades rotatable about the fan axis; and a pitch change mechanism operable with the plurality of fan blades, the pitch change mechanism including a plurality of linkages, the plurality of linkages including a first linkage coupled to a first fan blade of the plurality of fan blades and a second linkage coupled to a second fan blade of the plurality of fan blades; and a non-uniform blade actuator system operable with one or more of the plurality of linkages to control a pitch of the first fan blade relative to a pitch of the second fan blade.

Abstract: In one exemplary embodiment of the present disclosure an airfoil is provided. The airfoil defines a spanwise direction, a root end, and a tip end. The airfoil includes: a body defining a pressure side and a suction side and extending along the spanwise direction between the root end and the tip end, the body formed of a composite material; and a spar enclosed in the body of the airfoil extending along the spanwise direction, the spar comprising a plurality of spar segments arranged in an overlapping configuration along the spanwise direction.

Abstract: A rotor blade for a gas turbine engine is provided. The rotor blade includes a blade body formed of a first material; and a spar within a portion of the blade body, the spar formed of a second material that is different than the first material, the spar having an elongate body including a notch. The notch, weakened geometric feature, or other reduction in cross-section defines a frangible portion of the spar that is used to control a fracture of a rotor blade.

Abstract: A rotor blade for a gas turbine engine is provided. The rotor blade includes a blade body formed of a first material; and a tip component removably connected to the blade body, the tip component formed of a second material that is different than the first material.

Abstract: Methods, apparatus, systems, and articles of manufacture are disclosed for a slot to accommodate a blade of an engine. An example blade apparatus includes: a collet including a slot to receive an end of a blade; and a plurality of retainers including a first retainer positioned on a first end of the collet and a second retainer positioned on a second end of the collet to provide axial retention for the blade in the slot.

Abstract: An airfoil including a plurality of composite plies extending from a leading edge to a trailing edge and between a tip and a root. The airfoil further includes a frangible airfoil portion at the tip extending between the leading edge and the trailing edge and extending between the tip and a frangible line along a span. The frangible airfoil portion including a plurality of composite plies and one or more shape memory alloy inserts disposed between the plurality of composite plies. A gas turbine engine including a frangible airfoil and methods for forming a frangible airfoil are also disclosed.

Abstract: An aircraft defining a longitudinal direction and a lateral direction is provided. The aircraft includes: a fuselage; an engine mounted at a location spaced from the fuselage of the aircraft, the engine comprising a plurality of rotor blades; and a fuselage shield attached to or formed integrally with the fuselage at a location in alignment with the plurality of rotor blades along the lateral direction, the fuselage shield comprising a first layer defining a first density and a second layer defining a second density, the first density being different than the second density.

Abstract: An airfoil assembly comprising a central spar extending along the radial direction; a blade skin positioned around the central spar to define an airfoil that has a pressure side and a suction side extending in the axial direction between a leading edge and a trailing edge, wherein at least one cavity is defined between the blade skin and the central spar; and a support structure positioned at least partially within the at least one cavity. The support structure comprises: a foam; and a foam reinforcement structure embedded within the foam.

Abstract: Methods, apparatus, systems and articles of manufacture corresponding to a fan blade with intrinsic damping characteristics are disclosed. An example fan blade comprises an exterior body including a first side and a second side; a first hairpin structure in contact with (a) the first side of the exterior body and (b) the second side of the exterior body; and a second hairpin structure in contact with (a) the first side and (b) the second side, wherein the first hairpin structure corresponds to a first stiffness and the second hairpin structure corresponds to a second stiffness different than the first stiffness.

Abstract: An airfoil assembly comprising a central spar extending along the radial direction; a blade skin positioned around the central spar to define an airfoil that has a pressure side and a suction side extending in the axial direction between a leading edge and a trailing edge, wherein at least one cavity is defined between the blade skin and the central spar; and a support structure positioned at least partially within the at least one cavity. The support structure comprises: a foam; and a foam reinforcement structure embedded within the foam.

Abstract: A component for a turbine engine is provided. The component includes a main body portion having a flow path surface; and a protective layer formed of a chopped fiber material, the protective layer cohered to at least a portion of the flow path surface of the main body portion.

Abstract: A gas turbine engine is provided. The gas turbine engine includes: a turbomachine; a fan including a plurality of fan blades rotatably driven by the turbomachine; a nacelle surrounding at least in part the plurality of fan blades of the fan; and a clearance control system including a control ring positioned at least partially within the nacelle, coupled to the nacelle, or both for control of a clearance gap between the plurality of fan blades and the nacelle.

Abstract: An airfoil for gas turbine engines is generally provided. The airfoil defines a span extending between a root and a tip, the airfoil further defines a chord at each point along the span extending between a leading edge and a trailing edge. Further, the airfoil includes at least one frangible airfoil portion, a residual airfoil portion adjacent to the at least one frangible airfoil portion, and one or more zones. The one or more zones may include a plurality of cavities, where at least one or more cavities of the plurality of cavities include inclusions, with the inclusions having one or more materials.

Abstract: Composite components, such as gas turbine engine airfoils, and methods of forming composite components are provided. For example, a composite component of a gas turbine engine comprises a plurality of continuous fiber layers, each continuous fiber layer formed from a plurality of continuous fibers disposed in a matrix material, and a chopped fiber layer comprising a plurality of chopped fibers formed as a filmed sheet. The chopped fiber layer is laid up with the plurality of continuous fiber layers to form the composite component. A method for forming a composite component comprises laying up a plurality of continuous fiber layers, each continuous fiber layer comprising continuous fibers disposed in a matrix material; laying up a chopped fiber layer with the plurality of continuous fiber layers, the layup of the plurality of continuous fiber layers and the chopped fiber layer forming a reinforced layup; and curing the reinforced layup.

Abstract: A component for a gas turbine engine. The component includes a continuous fiber blade including an airfoil extending radially between a root and a tip and a blade attachment feature positioned at or adjacent to the root. The component further includes a platform coupled to the root of the continuous fiber blade. The platform includes a plurality of chopped fibers. Additionally, the component includes a thermoplastic polymer contained in both the continuous fiber blade and the platform. Moreover, the continuous fiber blade and platform are coupled together such that the continuous fiber blade and platform form a monolithic composite body.