A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. A speed reduction device such as an epicyclical gear assembly may be utilized to drive the fan section such that the fan section may rotate at a speed different than the turbine section so as to increase the overall propulsive efficiency of the engine. In such engine architectures, a shaft driven by one of the turbine sections provides an input to the epicyclical gear assembly that drives the fan section at a speed different than the turbine section such that both the turbine section and the fan section can rotate at closer to optimal speeds providing increased performance attributes and performance by desirable combinations of the disclosed features of the various components of the described and disclosed gas turbine engine.

Geared turbofan gas turbine engine architecture

A gas turbine engine includes a fan section, a gear arrangement configured to drive the fan section, a compressor section and a turbine section. The compressor section includes a low pressure compressor section and a high pressure compressor section. The turbine section is configured to drive compressor section and the gear arrangement. An overall pressure ratio, which is provided by a combination of a pressure ratio across said low pressure compressor section and a pressure ratio across said high pressure compressor section, is greater than about 35. The pressure ratio across the low pressure compressor section is between about 3 and about 8 whereas the pressure ratio across the high pressure compressor section is between about 7 and about 15.

Gas turbine engine compressor arrangement

A gas turbine engine system includes a nozzle having a plurality of positions for altering a discharge flow received through the nozzle from a gas turbine engine fan bypass passage. The nozzle is integrated with a thrust reverser having a stowed position and a deployed position to divert the discharge flow and generate a reverse thrust force. At least one actuator is coupled with the nozzle and the thrust reverser to selectively move the nozzle between the plurality of positions and to move the thrust reverser between the stowed position and the deployed position.

Dual function cascade integrated variable area fan nozzle and thrust reverser

A gas turbine engine according to an exemplary aspect of the present disclosure includes a gear train defined along an engine centeriine axis, and a spool along said engine centeriine axis which drives the gear train, the spool includes a low stage count low pressure turbine. In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the low stage count may include three to six (3-6) stages. Additionally or alternatively, the low stage count may include three (3) stages. Additionally or alternatively, the low stage count may include five (5) or six (6) stages.

Gas turbine engine with low stage count low pressure turbine

A fan drive gear system for a gas turbine engine includes a gear system that provides a speed reduction between a fan drive turbine and a fan and a mount flexibly supporting portions of the gear system A lubrication system supporting the fan drive gear system provides lubricant to the gear system and removes thermal energy produced by the gear system The lubrication system includes a capacity for removing thermal energy equal to less than about 2% of power input into the gear system

Fundamental gear system architecture

A turbofan engine control system is provided for managing a fan operating line. The engine (10) includes a spool having a turbine housed in a core nacelle (12). A turbofan (20) is coupled to the spool (14). A fan nacelle (34) surrounds the turbofan and core nacelle and provides a bypass flow path having a nozzle exit area (40). A controller (50) is programmed to effectively change the nozzle exit area in response to an undesired turbofan stability margin which may result in a stall or flutter condition. In one example, the physical nozzle exit area is increased at the undesired stability condition in which the airflow into the engine creates a destabilizing pressure gradient at the inlet side of the turbofan. A turbofan pressure ratio, turbofan pressure gradient, low spool speed and throttle position are monitored to determine the undesired turbofan stability margin.

Wayne Hurwitz

Papers

Method and device to avoid turbofan instability in a gas turbine engine