Gas turbine engine with fan variable area nozzle for low fan pressure ratio

Abstract

A gas turbine engine includes a core nacelle defined about an engine centerline axis. A core engine is at least partially disposed within the core nacelle. The core engine includes a first turbine section driving a first compressor section and a second turbine section driving a second compressor section. A fan section with a plurality of fan blades. A gear system is driven through a shaft coupled to the first turbine section. The gear system provides a speed reduction between the first turbine section and the fan section of greater than 2.3. A fan nacelle is mounted at least partially around said fan section and the core nacelle to define a fan bypass flow path for a fan bypass airflow. The fan bypass airflow has a fan pressure ratio of the fan bypass airflow during engine operation less than 1.45. A variable area fan nozzle is disposed at a trailing edge of the fan nacelle. The variable area fan nozzle includes a first nacelle section and a second nacelle section. The second nacelle section is movable relative to the first nacelle section to vary an exit area at the trailing edge of the fan nacelle to maintain the fan pressure ratio during engine operation.

What is claimed is: 
     
         1 . A gas turbine engine comprising:
 a core nacelle defined about an engine centerline axis;   a core engine at least partially disposed within the core nacelle, the core engine including a first turbine section driving a first compressor section and a second turbine section driving a second compressor section;   a fan section with a plurality of fan blades;   a gear system driven through a shaft coupled to the first turbine section, the gear system providing a speed reduction between the first turbine section and the fan section of greater than 2.3;   a fan nacelle mounted at least partially around said fan section and the core nacelle to define a fan bypass flow path for a fan bypass airflow, the fan bypass airflow having a fan pressure ratio of the fan bypass airflow during engine operation less than 1.45; and   a variable area fan bypass duct opening, the variable area fan bypass duct opening including a first nacelle section and a second nacelle section, the second nacelle section movable relative to the first nacelle section to vary an exit area from the fan stream to selectively lower the fan pressure ratio during engine operation.   
     
     
         2 . The engine as recited in  claim 1 , further comprising a controller operable to reduce said fan nozzle exit area at a cruise flight condition. 
     
     
         3 . The engine as recited in  claim 2 , wherein said controller is operable to control said fan nozzle exit area to reduce a fan instability. 
     
     
         4 . The engine as recited in  claim 2 , wherein said second nacelle section is movable axially substantially parallel to the engine centerline axis. 
     
     
         5 . The engine as recited in  claim 1 , wherein said fan section operates during engine operation at a corrected fan tip speed less than about 1150 ft/second. 
     
     
         6 . The engine as recited in  claim 5 , wherein said first turbine section operates at a pressure ratio that is greater than about five (5). 
     
     
         7 . The engine as recited in  claim 6 , wherein said gear system defines a gear reduction ratio of greater than or equal to about 2.5. 
     
     
         8 . The engine as recited in  claim 7 , wherein said fan bypass airflow defines a bypass ratio greater than about six (6). 
     
     
         9 . The engine as recited in  claim 8 , wherein said fan bypass airflow defines a bypass ratio greater than about ten (10). 
     
     
         10 . The engine as recited in  claim 9 , wherein the plurality of fan blades comprises twenty (20) or less fan blades. 
     
     
         11 . The engine as recited in  claim 1 , including an auxiliary port defined between the second nacelle section and the first nacelle section when the second nacelle section is moved axially aft, the auxiliary port defining a flow area for increasing bypass flow. 
     
     
         12 . The engine as recited in  claim 11 , including an entrance angle into the auxiliary port of less than twenty (20) degrees. 
     
     
         13 . The engine as recited in  claim 11 , wherein a flow area of the bypass flow path upstream of the auxiliary port is greater than a flow area through the auxiliary port. 
     
     
         14 . The engine as recited in  claim 13 , wherein the auxiliary port includes a forward most edge on in interior surface of the fan nacelle, wherein the forward most edge is spaced axially forward of a maximum radial distance between an inner surface of the second nacelle section and the engine centerline. 
     
     
         15 . The engine as recited in  claim 14 , wherein a duct cross-sectional area forward of the auxiliary port is greater than the port opening cross-sectional area. 
     
     
         16 . The engine as recited in  claim 15 , wherein said first turbine section operates at a pressure ratio that is greater than about five (5). 
     
     
         17 . The engine as recited in  claim 16 , wherein the second nacelle section is supported on a track faring for movement axially relative to the first nacelle section. 
     
     
         18 . The engine as recited in  claim 17 , wherein a cross-sectional area forward of the auxiliary port is greater than a cross-sectional area of the auxiliary port. 
     
     
         19 . The engine as recited in  claim 1 , including a plurality of fan exit guide vanes extending radially between the core nacelle and the fan nacelle aft of the fan section. 
     
     
         20 . The engine as recited in  claim 1 , wherein the second nacelle section comprises at least two sectors, with each of the two sectors movable independently.