Abstract: Gas turbine engine systems and related methods involving vane-blade count ratios greater than unity are provided. In this regard, a representative turbine stage for a gas turbine engine includes a first set of rotatable blades operative to be positioned downstream of and adjacent to a first set of vanes, a number of blades of the first set of blades being less than a number of vanes of the first set of vanes.

Abstract: A variable shape inlet guide vane (IGV) system includes a variable-shape IGV flap with a flexible portion that enables the desired spanwise distribution of Cx, alpha, and beta at a fan rotor inlet. An actuation system that rotates a root section of the variable-shape IGV flap to flex the flexible portion such that the twisted shape of the flap can reverse rather symmetrically during actuation from max open to max closed.

Abstract: A low pressure turbine for a gas turbine engine includes inner and outer counter-rotating rotor sets, with both said rotor sets driving a common shaft.

Abstract: Gas turbine engine systems and related methods involving vane-blade count ratios greater than unity are provided. In this regard, a representative turbine stage for a gas turbine engine includes a first set of rotatable blades operative to be positioned downstream of and adjacent to a first set of vanes, a number of blades of the first set of blades being less than a number of vanes of the first set of vanes.

Abstract: A variable shape inlet guide vane (IGV) system includes a variable-shape IGV flap with a flexible portion that enables the desired spanwise distribution of Cx, alpha, and beta at a fan rotor inlet. An actuation system that rotates a root section of the variable-shape IGV flap to flex the flexible portion such that the twisted shape of the flap can reverse rather symmetrically during actuation from max open to max closed.

Abstract: A turbine blade is cooled by cooling air that flows through a radial cooling channel. The turbine blade includes a root and an airfoil. The flow of cooling air into the cooling channel is limited by a pre-meter orifice to provide a reduced pressure within the cooling channel. The pressure drop results from the cross-sectional area of the pre-meter orifice being less than the cross-sectional area of the adjacent cooling channel. After flowing through the cooling channel, the cooling air exits the cooling channel through a film hole to form a film layer over the airfoil to cool and insulate the turbine blade.

Abstract: A rotor blade for a gas turbine engine includes a bowed surface on a tip region of the suction side thereof. The curvature of the bowed surface progressively increases toward the tip of the blade. The bowed surface results in a reduction of tip leakage through a tip clearance from the pressure side to the suction side of the blade and reduces mixing loss due to tip leakage.

Abstract: The first stage of vanes (16) and second stage of vanes (24) each contain the same number of vanes. The second stage vanes are located such that the wake flow (38) from the first stage vanes falls on or near the leading edge, after passing through the stage of rotating blades.