Gas turbine engine with low stage count low pressure turbine

Abstract

A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a spool along an engine axis which drives a gear train, the spool including a low stage count low pressure turbine.

What is claimed is: 
     
         1 . A gas turbine engine comprising:
 a gear train positioned along an engine axis;   a first turbine section configured for driving said gear train, said first turbine section including between three to six (3-6) stages;   a second turbine section including at least two (2) stages;   a fan including a plurality of fan blades rotatable about the engine axis and driven by at least one of the first turbine section and the second turbine section through the gear train;   a core nacelle mounted about the engine axis; and   a fan nacelle mounted about the core nacelle to define a bypass flow path,   
       wherein a bypass ratio defined by the bypass flow path airflow divided by airflow through the core nacelle is greater than about six (6). 
     
     
         2 . The engine as recited in  claim 1 , wherein said first turbine section includes three (3) or four (4) stages. 
     
     
         3 . The engine as recited in  claim 1 , wherein said first turbine section includes five (5) stages. 
     
     
         4 . The engine as recited in  claim 1 , wherein said first turbine section includes six (6) stages. 
     
     
         5 . The engine as recited in  claim 1 , further including
 a fan variable area nozzle configured to vary a fan nozzle exit area and adjust a pressure ratio of the fan bypass airflow during engine operation.   
     
     
         6 . The engine as recited in  claim 5 , further comprising:
 a controller operable to control the fan variable area nozzle to vary the fan nozzle exit area and adjust the pressure ratio of the fan bypass airflow.   
     
     
         7 . The engine as recited in  claim 1 , wherein said gear train defines a gear reduction ratio of greater than or equal to about 2.5. 
     
     
         8 . The engine as recited in  claim 7 , wherein said first turbine section defines a pressure ratio that is greater than about five (5.0). 
     
     
         9 . The engine as recited in  claim 1 , wherein a fan pressure ratio across the fan blades is less than about 1.45. 
     
     
         10 . The engine as recited in  claim 9 , wherein the bypass ratio is between about six and about ten (6.0-10.0). 
     
     
         11 . The engine as recited in  claim 1 , wherein the first turbine section is configured to drive the fan through the gear train to provide a low corrected fan tip speed of less than about 1150 feet/second. 
     
     
         12 . The engine as recited in  claim 1 , wherein the bypass ratio is between about six and about ten (6.0-10.0). 
     
     
         13 . The engine as recited in  claim 1 , wherein said first turbine section is one of three turbine rotors, while said second turbine section and another one of said turbine rotors each drives a compressor section. 
     
     
         14 . The engine as recited in  claim 13 , wherein said first turbine section drives a compressor section. 
     
     
         15 . The engine as recited in  claim 14 , wherein said gear train is positioned intermediate a compressor section driven by said first turbine section and said fan. 
     
     
         16 . The engine as recited in  claim 14 , wherein said gear train is positioned intermediate said first turbine section and said compressor section driven by said first turbine section. 
     
     
         17 . A method of designing a gas turbine engine comprising:
 providing a core nacelle defined about an engine axis;   providing a fan nacelle mounted at least partially around said core nacelle to define a fan bypass flow path for a fan bypass airflow;   providing a gear train within said core nacelle;   providing a first spool along said engine axis within said core nacelle to drive said gear train, said first spool includes a first turbine section including between three-six ( 3 - 6 ) stages, and a first compressor section;   providing a second spool along said engine axis within said core nacelle, said second spool includes a second turbine section including at least two ( 2 ) stages and a second compressor section; and   providing a fan including a plurality of fan blades to be driven through the gear train by the first spool, wherein the bypass flow path is configured to provide a bypass ratio of airflow through the bypass flow path divided by airflow through the core nacelle that is greater than about six (6) during engine operation.   
     
     
         18 . The method as recited in  claim 17 , wherein said first turbine section defines a pressure ratio that is greater than about five (5.0). 
     
     
         19 . The method as recited in  claim 18 , wherein a fan pressure ratio across the plurality of fan blades is less than about 1.45. 
     
     
         20 . The method as recited in  claim 19 , wherein the gear train is configured to provide a speed reduction ratio greater than about 2.5:1. 
     
     
         21 . The method as recited in  claim 20 , wherein the plurality of fan blades are configured to rotate at a fan tip speed of less than about 1150 feet/second during engine operation. 
     
     
         22 . The method as recited in  claim 21 , wherein the second turbine section includes two (2) stages. 
     
     
         23 . The method as recited in  claim 22 , wherein the first turbine section includes three (3) turbine stages. 
     
     
         24 . The method as recited in  claim 23 , wherein the bypass ratio is between about six and about ten (6.0-10.0).