Abstract:
Aspects of the disclosure are directed to a gas turbine engine, comprising a sealing assembly that includes a non-contacting HALO seal, a ring constructed as a full ring and configured as a carrier of the seal, a sealing land configured to rotate and interface to the seal. In some embodiments, the engine further comprises a vane, where the ring is coupled to the vane.
Abstract:
A three-spool turbofan engine (20) has a variable fan nozzle (35). The fan blades have a peak tip radius RT and an inboard leading edge radius RH at an inboard boundary of the flowpath. A ratio of RH to RT is less than about 0.40.
Abstract:
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.
Abstract:
A gas turbine engine comprises a fan drive turbine. The fan drive turbine drives the fan through a gear reduction; A change in enthalpy is defined across the gas turbine engine. The change in enthalpy divided by a speed of the fan drive turbine squared is less than or equal to about 1.8. An axial component of gases approaching an upstream most blade of the fan drive turbine divided by the speed of the fan drive turbine is equal to or less than about 0.9.
Abstract:
A fan section for use in a gas turbine engine has a fan rotor with a plurality of blades and an outer fan housing surrounding the plurality of blades. A tip clearance is defined between a radially outer tip of the blades and a radially inner surface of the fan housing. A fan drive shaft drives the rotor. A drive input drives the fan drive shaft. A shifting mechanism shifts a location of the blades relative to the drive input, thereby controlling the tip clearance. A gas turbine engine is also disclosed.
Abstract:
A front section for a gas turbine engine according to an example of the present disclosure includes, among other things, a fan section including a fan hub. The fan hub includes a hub diameter supporting a plurality of fan blades including a tip diameter. A transitional entrance passage is configured to communicate flow between the fan section and a compressor section. The transitional entrance passage includes an inlet disposed at an inlet diameter and an outlet to the compressor section disposed at an outlet diameter. A method of designing a gas turbine engine is also disclosed.
Abstract:
A gas turbine engine according to an example of the present disclosure includes, among other things, a fan section, and a low spool including a low pressure compressor section and a low pressure turbine. A high spool includes a high pressure compressor section. A gear arrangement is defined along an engine axis. The low spool is operable to drive the fan section through the gear arrangement. A mount system includes an aft mount configured to react at least a portion of a thrust load at an engine case generally parallel to an engine axis.
Abstract:
A turbine section of a gas turbine engine according to an example of the present disclosure includes, among other things, a first turbine section and a second turbine section. The first turbine section has a first exit area and rotates at a first speed. The second turbine section has a second exit area and rotates at a second speed. A first performance quantity is defined as the product of the first speed squared and the first exit area. A second performance quantity is defined as the product of the second speed squared and the second exit area.
Abstract:
A gas turbine engine comprises at least two compressor rotors, including a first lower pressure compressor rotor and a second higher pressure compressor rotor. At least two corresponding air taps include a low tap for tapping low pressure compressor air from a location downstream of a first stage of the lower pressure compressor rotor, and upstream of a first stage of the higher pressure compressor rotor, and a high tap to tap air downstream of the first stage of the higher pressure compressor rotor. an air handling system selectively communicates both the low tap and the high tap to an air use destination. Air is selectively supplied from the low tap to the air handling system at a high power operation and from the high tap to the air handling system at a low power operation.
Abstract:
A turbine section for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a fan drive turbine including a fan drive duct, the fan drive turbine being configured to drive a fan section through a geared architecture at a speed that is less than an input speed to the geared architecture. A fan drive turbine includes a fan drive duct, the fan drive turbine being configured to drive a fan section through a geared architecture at a speed that is less than an input speed to the geared architecture. At least one upstream turbine is configured to drive at least one compressor. The at least one upstream turbine includes a turbine duct defining a conical flow path having a conical inlet defined by a first diameter and a conical outlet defined by a second diameter greater than the first diameter. The conical outlet is in fluid communication with the fan drive duct downstream of the conical outlet. At least one row of shrouded rotor blades defines at least a portion of the conical flow path. A method of designing a gas turbine engine is also disclosed.