A gas turbine engine that has improved fuel burn provides operability and/or maintenance requirements when installed on an aircraft. The gas turbine engine is provided with a core compressor that includes twelve, thirteen or fourteen rotor stages. The gas turbine engine has a ratio of a core compressor aspect ratio divided by a core compressor pressure ratio is in the range of from 0.03 to 0.09. 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 variable geometry inlet system of an aircraft engine includes an inlet duct. The inlet duct includes at least first and second sections moveable between extended and retracted positions such that the inlet duct defines a variable axial length of an inlet passage for selective flight conditions. The inclusion of acoustic treatment may assist in controlling noise.

A propulsion system according to an example of the present disclosure includes, among other things, a geared architecture configured to drive a fan section including a fan, and a turbine section configured to drive the geared architecture. The turbine section has an exit point, and a diameter (Dt) defined as the outer diameter of a last blade airfoil stage in the turbine section at the exit point. A nacelle surrounds a core engine housing. The fan is configured to deliver air into a bypass duct defined between the nacelle and the core engine housing. A core engine exhaust nozzle is downstream of the exit point. A downstream most point of the core engine exhaust nozzle is defined at a distance (Lc or Ln) from the exit point.

A propulsion system according to an example of the present disclosure includes, among other things, a geared architecture configured to drive a fan section including a fan, and a turbine section configured to drive the geared architecture. The turbine section has an exit point, and a diameter (Dt) defined as the outer diameter of a last blade airfoil stage in the turbine section at the exit point. A nacelle surrounds a core engine housing. The fan is configured to deliver air into a bypass duct defined between the nacelle and the core engine housing. A core engine exhaust nozzle is downstream of the exit point. A downstream most point of the core engine exhaust nozzle is defined at a distance (Lc or Ln) from the exit point.

The present disclosure is directed to a gas turbine engine including a turbine section including a first rotating component interdigitated along a longitudinal direction with a second rotating component. The first rotating component and the second rotating component are each coupled to a speed reduction assembly in counter-rotating arrangement. The first rotating component comprising an outer shroud and a plurality of outer shroud airfoils extended inward along a radial direction from the outer shroud. A connecting member couples the outer shroud to a radially extended first rotor. The second rotating component comprising an inner shroud and a plurality of inner shroud airfoils extended outward along the radial direction from the inner shroud, the plurality of inner shroud airfoils in alternating arrangement along the longitudinal direction with the plurality of outer shroud airfoils. The gas turbine engine defines a radius per unit thrust defined by a maximum radius at the turbine section over a maximum thrust output between approximately 0.0004 to approximately 0.0010 inches per pound thrust.

A highly efficient gas turbine engine is a system wherein the fan of the gas turbine engine is driven from a turbine via a gearbox, such that the fan has a lower rotational speed than the driving turbine, thereby providing efficiency gains. The efficient fan system is mated to a core that has low cooling flow requirements and/or high temperature capability, and which may have particularly low mass for a given power.

A variable geometry inlet system of an aircraft engine includes an inlet duct. The inlet duct includes at least first and second sections moveable between extended and retracted positions such that the inlet duct defines a variable axial length of an inlet in controlling noise.

An aircraft gas turbine engine comprises a fan coupled to a fan drive turbine, the fan being configured to provide a bypass flow (B) and a core flow (A) in use. The engine includes a reduction gearbox which couples the fan to the fan drive turbine and a core compressor arrangement. The core compressor arrangement has a core inlet at an upstream end of a core gas flow passage (A) defined by radially inner and outer walls, and at least a first compressor rotor blade provided at an upstream end of the compressor arrangement. The radially inner wall of the core inlet defines a first diameter (D.sub.INLET), and a root leading edge of the first compressor rotor blade defines a second diameter (D.sub.COMP). A first ratio (D.sub.INLET:D.sub.COMP) of the first diameter (D.sub.COMP) to the second diameter (D.sub.COMP) is greater than or equal to 1.4.

An aircraft gas turbine engine comprises a fan coupled to a fan drive turbine, the fan being configured to provide a bypass flow (B) and a core flow (A) in use. The engine includes a reduction gearbox which couples the fan to the fan drive turbine and a core compressor arrangement. The core compressor arrangement has a core inlet at an upstream end of a core gas flow passage (A) defined by radially inner and outer walls, and at least a first compressor rotor blade provided at an upstream end of the compressor arrangement. The radially inner wall of the core inlet defines a first diameter (D.sub.INLET), and a root leading edge of the first compressor rotor blade defines a second diameter (D.sub.COMP). A first ratio (D.sub.INLET:D.sub.COMP) of the first diameter (D.sub.COMP) to the second diameter (D.sub.COMP) is greater than or equal to 1.4.

A flow control system on an aircraft engine nacelle incorporates a plurality of translating turning vanes each having a body. An equal plurality of actuators is coupled to a trailing edge of the body of an associated one of the translating turning vanes. The actuator translates the body from a retracted position to an extended position.