Abstract: According to an example embodiment, a gas turbine engine assembly includes, among other things, a fan that has a plurality of fan blades. A diameter of the fan has a dimension D that is based on a dimension of the fan blades. Each fan blade has a leading edge. An inlet portion is situated forward of the fan. A length of the inlet portion has a dimension L between a location of the leading edge of at least some of the fan blades and a forward edge on the inlet portion. A dimensional relationship of L/D is between about 0.2 and about 0.45.

Abstract: According to an example embodiment, a gas turbine engine assembly includes, among other things, a fan that has a plurality of fan blades. A diameter of the fan has a dimension D that is based on a dimension of the fan blades. Each fan blade has a leading edge. An inlet portion is situated forward of the fan. A length of the inlet portion has a dimension L between a location of the leading edge of at least some of the fan blades and a forward edge on the inlet portion. A dimensional relationship of L/D is between about 0.2 and 0.45.

Abstract: A nacelle assembly for a high-bypass gas turbine engine includes a core nacelle defined about an engine centerline axis. A fan nacelle is mounted at least partially around the core nacelle to define a fan bypass flow path. A variable area fan nozzle is in communication with the fan bypass flow path. The variable area fan nozzle has a first fan nacelle section and a second fan nacelle section. The second fan nacelle section is axially movable relative to the first fan nacelle section to define an auxiliary port at a non-closed position to vary a fan nozzle exit area and adjust fan bypass airflow. The second fan nacelle section includes an acoustic system that has an acoustic impedance located on a radially outer surface.

Abstract: A first resource is used for initiating a transaction on a first channel. A second resource of a second channel is identified to assist in a portion of the transaction. The second resource is claimed for the transaction and the portion of the transaction is routed to the second resource on the second channel for processing.

Abstract: A gas turbine engine assembly according to an example of the present disclosure includes, among other things, a 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, a geared architecture configured to drive the fan, a turbine section configured to drive the geared architecture, a compressor section including a first compressor and a second compressor, and an inlet portion forward of the fan. A length of the inlet portion has a dimension L between a location of the leading edge of at least some of the fan blades and a forward edge on the inlet portion. A dimensional relationship of L/D is between about 0.2 and about 0.45.

Abstract: A fan section for a gas turbine engine includes a fan hub and a nose cone section operatively mounted to the fan hub. The nose cone section includes a noise attenuation feature.

Abstract: A rotor blade comprises an airfoil extending radially from a root section to a tip section and axially from a leading edge to a trailing edge, the leading and trailing edges defining a curvature therebetween. The curvature determines a relative exit angle at a relative span height between the root section and the tip section, based on an incident flow velocity at the leading edge of the airfoil and a rotational velocity at the relative span height. In operation of the rotor blade, the relative exit angle determines a substantially flat exit pressure ratio profile for relative span heights from 75% to 95%, wherein the exit pressure ratio profile is constant within a tolerance of 10% of a maximum value of the exit pressure ratio profile.

Abstract: 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.

Abstract: A nacelle assembly for a high-bypass gas turbine engine includes a core nacelle defined about an engine centerline axis. A fan nacelle is mounted at least partially around the core nacelle to define a fan bypass flow path. A variable area fan nozzle is in communication with the fan bypass flow path. The variable area fan nozzle has a first fan nacelle section and a second fan nacelle section. The second fan nacelle section is axially movable relative to the first fan nacelle section to define an auxiliary port at a non-closed position to vary a fan nozzle exit area and adjust fan bypass airflow. The second fan nacelle section includes an acoustic system that has an acoustic impedance located on a radially outer surface.

Abstract: A bypass gas turbine engine includes a variable area fan nozzle with a second fan nacelle section axially movable relative to a first fan nacelle section to define an auxiliary port to vary a fan nozzle exit area and adjust fan bypass airflow. The second fan nacelle section includes a leading edge region that defines a concave external contour and the first fan nacelle section includes a trailing edge region that defines a convex internal contour.

Abstract: A high-bypass gas turbine engine includes a variable area fan nozzle with an acoustic system having an acoustic impedance.

Abstract: According to an example embodiment, a gas turbine engine assembly includes, among other things, a fan that has a plurality of fan blades. A diameter of the fan has a dimension D that is based on a dimension of the fan blades. Each fan blade has a leading edge. An inlet portion is situated forward of the fan. A length of the inlet portion has a dimension L between a location of the leading edge of at least some of the fan blades and a forward edge on the inlet portion. A dimensional relationship of L/D is between about 0.2 and about 0.45.

Abstract: A nacelle assembly according to an exemplary aspect of the present disclosure includes, among other things, a core nacelle defined at least partially about a core engine, a fan nacelle mounted at least partially around the core nacelle to define a fan bypass flow path, and a variable area fan nozzle in communication with the fan bypass flow path. The variable area fan nozzle has a first fan nacelle section and a second fan nacelle section downstream of the first fan nacelle section. The first fan nacelle section and the second fan nacelle section are axially movable relative to one another to define an auxiliary port to vary a fan nozzle exit area and adjust fan bypass airflow. The auxiliary port is defined between the first fan nacelle section and the second fan nacelle section. The first fan nacelle section comprises a first acoustic system which provides an acoustic impedance configured to attenuate a noise characterized by a leading edge of the second fan nacelle section.

Abstract: A high-bypass gas turbine engine includes a variable area fan nozzle with an acoustic system having an acoustic impedance.

Abstract: A bypass gas turbine engine includes a variable area fan nozzle with a leading edge region that defines an increased airfoil leading edge radius.

Abstract: A nacelle structure for a gas turbine engine assembly includes a fan case, an inlet, and a noise attenuation device. The fan case is configured to be disposed about a fan section of the gas turbine engine, which fan section has a diameter D. The inlet is attached to the fan case and extends axially forward of the fan case. A hilite of the inlet is spaced a distance L from a region in which the fan section is configured to be disposed. A ratio L/D is less than about 0.6. The noise attenuation structure covers a portion of an inner surface of the fan case and the inlet.

Abstract: A fan section for a gas turbine engine has a fan rotor with a plurality of fan blades. A plurality of exit guide vanes are positioned to be downstream of the fan rotor. The fan rotor is driven through a gear reduction relative to a turbine section. The exit guide vanes are desired to address resultant sound from interaction of wakes from the fan blades across exit guide vanes. A gas turbine engine incorporating a fan section is also disclosed.

Abstract: A bypass gas turbine engine includes a variable area fan nozzle with a second fan nacelle section axially movable relative to a first fan nacelle section to define an auxiliary port to vary a fan nozzle exit area and adjust fan bypass airflow. The second fan nacelle section includes a leading edge region that defines a concave external contour and the first fan nacelle section includes a trailing edge region that defines a convex internal contour.

Abstract: A high-bypass gas turbine engine includes a variable area fan nozzle with an acoustic system having an acoustic impedance.

Abstract: A bypass gas turbine engine includes a variable area fan nozzle with a leading edge region that defines an increased airfoil leading edge radius.