A gas turbine has a rotor blade. The rotor blade has a blade root connected to an airfoil. The blade root has a root contour with respect to a cross-sectional view. From a lower end of the blade root, the blade root contour has convex contour portions and concave contour portions. From the lower end along the blade root contour between a convex contour portion and an adjoining concave contour portion, there is a contour portion as a flank portion that is load-bearing. From the lower end along the blade root contour between a concave contour portion and an adjoining convex contour portion, there is a contour portion as a flank portion that is not load-bearing in operation. At least one of the concave contour portions has a first arc portion, a second arc portion, and a straight portion disposed between the two arc portions.

A gas turbine engine has a compression system blade ratio defined as the ratio of the height of a fan blade to the height of the most downstream compressor blade in the range of from 45 to 95. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

A gas turbine has a rotor blade. The rotor blade has a blade root connected to an airfoil. The blade root has a root contour with respect to a cross-sectional view. From a lower end of the blade root, the blade root contour has convex contour portions and concave contour portions. From the lower end along the blade root contour between a convex contour portion and an adjoining concave contour portion, there is a contour portion as a flank portion that is load-bearing. From the lower end along the blade root contour between a concave contour portion and an adjoining convex contour portion, there is a contour portion as a flank portion that is not load-bearing in operation. At least one of the concave contour portions has a first arc portion, a second arc portion, and a straight portion disposed between the two arc portions.

A geared turbofan engine includes a first rotor, a fan, a second rotor, a gear train, a fan casing, a nacelle and a plurality of discrete acoustic liner segments. The fan is connected to the first rotor and is capable of rotation at frequencies between 200 and 6000 Hz and has a fan pressure ratio of between 1.25 and 1.60. The gear train connects the first rotor to the second rotor. The fan casing and nacelle are arranged circumferentially about a centerline and define a bypass flow duct in which the fan is disposed. The plurality of discrete acoustic liner segments with varied geometric properties are disposed along the bypass flow duct.

According to an example embodiment, a gas turbine engine assembly includes, among other things, a fan section including a fan, the fan including a plurality of fan blades, a diameter of the fan having a dimension D that is based on a dimension of the fan blades, each fan blade having a leading edge, and a forward most portion on the leading edges of the fan blades in a first reference plane, a geared architecture, a turbine section including a high pressure turbine and a low pressure turbine, the low pressure turbine driving the fan through the geared architecture, a nacelle surrounding the fan, the nacelle including an inlet portion forward of the fan, a forward edge on the inlet portion in a second reference plane, and a length of the inlet portion having a dimension L measured along an engine axis between the first reference plane and the second reference plane. A dimensional relationship of L/D is between 0.20 and 0.40.

A nacelle assembly of a gas turbine engine includes an annular structure defining a central axis, and having a radially inward surface and a radially outward surface, the radially inward surface at least partially defining a bypass duct. An aft portion of the radially inward surface at least partially defines an axially extending convergent-divergent exit nozzle. A secondary nozzle flap is radially spaced from the aft portion of the radially inward surface. The secondary nozzle flap and the aft portion of the radially inward surface define a secondary bypass duct therebetween. The secondary nozzle flap is operably connected to the annular structure such that the secondary nozzle flap is selectably movable relative to the aft portion of the radially inward surface, thereby changing a cross-sectional area of a secondary bypass duct exit.

A gas turbine engine has a compression system blade ratio defined as the ratio of the height of a fan blade to the height of the most downstream compressor blade in the range of from 45 to 95. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

A gas turbine engine has a compression system blade ratio defined as the ratio of the height of a fan blade to the height of the most downstream compressor blade in the range of from 45 to 95. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

A blade of a turbo machine, having a blade leaf, with a flow leading edge, a flow trailing edge, and flow conducting surfaces, and a cooling passage integrated in the blade leaf. In the region of the blade leaf cooling passage portions extend substantially in the radial direction. Adjacent cooling passage portions merge into one another via a diversion passage portion having a material web extending between the adjacent cooling passage portions. The respective material web ends in the region of the respective diversion passage portion. The respective material web has a defined axial width between the respective adjacent cooling passage portions and the respective material web in the region of the respective diversion passage portion has a material thickening enlarging the axial width by at least 20%.

A gas turbine engine comprises a fan rotor having a hub and a plurality of fan blades extending radially outwardly of the hub. A compressor is positioned downstream of the fan rotor, and has a first compressor blade row defined along a rotational axis of the fan rotor and the compressor rotor. A gear reduction is positioned axially between the first compressor blade row and the fan rotor, and includes a ring gear and a carrier. The carrier has an axial length and the ring gear has an outer diameter. A ratio of the axial length to the outer diameter may be greater than or equal to about 0.20 and less than or equal to about 0.40. The gear reduction is connected to drive the hub to rotate. A method of designing a gas turbine engine is also disclosed.