An array of flow-directing elements for a compressor of a gas turbine including at least one first flow-directing element and at least one second flow-directing element different from the first flow-directing element; the flow-directing elements each having a leading edge facing the gas turbine inlet, a trailing edge, a pressure side connecting them and located ahead in the direction of rotation, a suction side located opposite thereof, as well as successive chords along a stacking axis; the flow-directing elements each extending between an airfoil root proximate to the rotor and an airfoil tip remote from the rotor. The trailing edge of the first flow-directing element is, at least in a portion thereof, axially offset from the trailing edge of the second flow-directing element in a direction toward the leading edge at least in a half proximate to the airfoil tip.

An array of flow-directing elements for a compressor of a gas turbine including at least one first flow-directing element and at least one second flow-directing element different from the first flow-directing element; the flow-directing elements each having a leading edge facing the gas turbine inlet, a trailing edge, a pressure side connecting them and located ahead in the direction of rotation, a suction side located opposite thereof, as well as successive chords along a stacking axis; the flow-directing elements each extending between an airfoil root proximate to the rotor and an airfoil tip remote from the rotor. The trailing edge of the first flow-directing element is, at least in a portion thereof, axially offset from the trailing edge of the second flow-directing element in a direction toward the leading edge at least in a half proximate to the airfoil tip.

An assembly for a gas turbine engine according to an aspect of the present disclosure includes a metallic damper including a first contact surface and a gas turbine engine component. The gas turbine engine component includes a main body extending in a first direction between a gaspath surface and a second contact surface. The first and second contact surfaces oppose each other along an interface extending in a second direction. The first and second contact surfaces are dimensioned to contact each other along the interface in a hot assembly state. The main body is established by a composite including fibers in a matrix material. At least some of the fibers are arranged to establish a plurality of cooling passages aligned with the interface relative to the second direction. A method of damping for a gas turbine engine is also disclosed.

Airfoil vibration damping apparatus are disclosed. An example apparatus includes a metallic airfoil including a cavity, and a dilatant material disposed in the cavity to dampen vibrations of the metallic airfoil.

Airfoil vibration damping apparatus are disclosed. An example apparatus includes a metallic airfoil including a cavity, and a dilatant material disposed in the cavity to dampen vibrations of the metallic airfoil.

A blisk fan is provided for a turbine engine propulsion system. The blisk fan includes a hub configured to rotate about a rotational axis at a maximum rotational speed, and a plurality of blades extending radially outward from the hub to define a fan leading edge tip diameter. Each of the blades has a first vibratory mode at a natural frequency, which is greater than a first fan order and less than a second fan order at the maximum rotational speed. The compression system preferably has a balance factor of the compression system between 1.9 and 3.2.

A blade assembly may be configured to rotate about an axis of a gas turbine engine. The blade assembly may include a blade and a blade cover. A blade may include a first surface defining a perimeter of a cavity in the blade. A first portion of the perimeter of the cavity includes a lip. A blade cover disposed over the cavity in the blade, such that the cavity is disposed between the blade and the blade cover, and under the lip, such that blade cover is between a portion of the blade and the cavity.

A turbomachine blade assembly including a turbomachine blade (1), in particular for a gas turbine, and at least one tuning element container including a housing (10) attached to the turbomachine blade and an insert (20) disposed in a recess (11) of this housing. A wall (20; 21) of the insert spaces apart two first cavities (31), which each accommodate at least one tuning element (40) provided for impacting contact with the housing (10) and the insert (20).

An airfoiled component comprises: a root section, an airfoil section, a damper pocket enclosed within a portion of the airfoil section, and a damper. The airfoil section includes a suction sidewall and a pressure sidewall each extending chordwise between a leading edge and a trailing edge, and extending spanwise between the root section and an airfoil tip. The damper includes a fixed end unified with a damper mounting surface, and a free end extending into the damper pocket from the damper mounting surface.

An airfoiled component comprises: a root section, an airfoil section, a damper pocket enclosed within a portion of the airfoil section, and a damper. The airfoil section includes a suction sidewall and a pressure sidewall each extending chordwise between a leading edge and a trailing edge, and extending spanwise between the root section and an airfoil tip. The damper includes a fixed end unified with a damper mounting surface, and a free end extending into the damper pocket from the damper mounting surface.