Method for setting a gear ratio of a fan drive gear system of a gas turbine engine

Abstract

A gas turbine engine includes an engine centerline longitudinal axis and a fan section including a fan with fan blades and rotatable about the engine centerline longitudinal axis. A low corrected fan tip speed less than about 1400 ft/sec and the low corrected fan tip speed is an actual fan tip speed determined at an ambient temperature divided by [(Tram° R)/( 518.7 ° R)] 0.5 , where T represents the ambient temperature in degrees Rankine. A bypass ratio greater than about 11 and a speed reduction device having a gear system with a gear ratio and a plurality of bearing systems. A low and high speed spool including a low and high pressure turbine and a first and second shaft, respectively. The first and second shafts are concentric and mounted via at least one of the bearing systems for rotation about the engine centerline longitudinal axis and the first shaft is in communication with the fan through the speed reduction device and the low pressure turbine includes four stages.

What is claimed is: 
     
         1 . A gas turbine engine comprising:
 an engine centerline longitudinal axis;   a fan section including a fan with a plurality of fan blades and rotatable about the engine centerline longitudinal axis, wherein the fan has a low corrected fan tip speed less than about  1400  ft/sec, wherein the low corrected fan tip speed is an actual fan tip speed at an ambient temperature divided by [(Tram° R)/(518.7° R)] 0.5 , where T represents the ambient temperature in degrees Rankine;   a bypass ratio greater than about 11.0;   a fan pressure ratio less than 1.48, wherein the fan pressure ratio is measured across a fan blade alone;   a speed reduction device comprising a gear system with a gear ratio of at least 2.6 and less than or equal to 4.1;   a plurality of bearing systems;   a low speed spool comprising a low pressure turbine and a first shaft; and   a high speed spool comprising a high pressure turbine and a second shaft;   wherein the first shaft and second shaft are concentric and mounted via at least one of the plurality of bearing systems for rotation about the engine centerline longitudinal axis, and the first shaft is in communication with the fan through the speed reduction device;   wherein the low pressure turbine includes at least three stages and no more than four stages and a pressure ratio greater than about 5:1, the low pressure turbine includes an inlet having an inlet pressure, and an outlet that is prior to any exhaust nozzle and having an outlet pressure, and the pressure ratio of the low pressure turbine is a ratio of the inlet pressure to the outlet pressure; and   wherein the high pressure turbine includes two stages.   
     
     
         2 . The gas turbine engine of  claim 1 , wherein the bypass ratio is greater than 13 and less than 20. 
     
     
         3 . The gas turbine engine of  claim 2 , wherein the low pressure turbine includes four stages. 
     
     
         4 . The gas turbine engine of  claim 3 , wherein the gear system is a star gear system with a ring gear, and a sun gear, wherein the gear ratio is determined by measuring a diameter of the ring gear and dividing that diameter by the diameter of the sun gear. 
     
     
         5 . The gas turbine engine of  claim 4 , wherein the low pressure turbine includes at least one rotor constrained by a first stress level, at least one of the plurality of fan blades of the fan constrained by a second stress level and having a fan tip speed boundary condition, and the gear ratio is configured such that in operation the fan blade does not exceed the fan tip speed boundary condition or the second stress level, and the low pressure turbine rotor does not exceed the first stress level. 
     
     
         6 . The gas turbine engine of  claim 4 , wherein the fan pressure ratio is less than 1.38. 
     
     
         7 . A gas turbine engine comprising:
 an engine centerline longitudinal axis;   a fan section including a fan with a plurality of fan blades and rotatable about the engine centerline longitudinal axis;   a bypass ratio greater than about 11.0;   a fan pressure ratio less than 1.52, wherein the fan pressure ratio is measured across a fan blade alone;   a speed reduction device comprising a gear system;   a plurality of bearing systems;   a low speed spool comprising a low pressure turbine and a first shaft; and   a high speed spool comprising a high pressure turbine and a second shaft;   wherein the first shaft and second shaft are concentric and mounted via at least one of the plurality of bearing systems for rotation about the engine centerline longitudinal axis, and the first shaft is in communication with the fan through the speed reduction device;   wherein the low pressure turbine has a pressure ratio greater than about 5:1, the low pressure turbine includes an inlet having an inlet pressure, and an outlet that is prior to any exhaust nozzle and having an outlet pressure, and the pressure ratio of the low pressure turbine is a ratio of the inlet pressure to the outlet pressure; and   wherein the high pressure turbine includes two stages.   
     
     
         8 . The gas turbine engine of  claim 7 , wherein the low pressure turbine includes five stages. 
     
     
         9 . The gas turbine engine of  claim 7 , further comprising a low pressure compressor including three stages, wherein the low pressure turbine is configured to drive the low pressure compressor. 
     
     
         10 . The gas turbine engine of  claim 7 , wherein the low pressure turbine includes at least three stages and no more than four stages. 
     
     
         11 . The gas turbine engine of  claim 10 , wherein the low pressure turbine includes four stages. 
     
     
         12 . The gas turbine engine of  claim 11 , further comprising a low corrected fan tip speed less than about 1400 ft/sec, wherein the low corrected fan tip speed is an actual fan tip speed at an ambient temperature divided by [(Tram° R)/(518.7° R)] 0.5 , where T represents the ambient temperature in degrees Rankine. 
     
     
         13 . The gas turbine engine of  claim 12 , further comprising a low pressure compressor including three stages, wherein the low pressure turbine is configured to drive the low pressure compressor. 
     
     
         14 . The gas turbine engine of  claim 13 , further comprising a high pressure compressor including eight stages, wherein the high pressure turbine is configured to drive the high pressure compressor. 
     
     
         15 . The gas turbine engine of  claim 12 , wherein the gear system is a star gear system with a ring gear, a sun gear, and a star gear ratio, and the star gear ratio is determined by measuring a diameter of the ring gear and dividing that diameter by the diameter of the sun gear. 
     
     
         16 . The gas turbine engine of  claim 15 , the star gear system further comprising five intermediate gears. 
     
     
         17 . The gas turbine engine of  claim 16 , further comprising a mid-turbine frame arranged between the high pressure turbine and the low pressure turbine, wherein the mid-turbine frame supports at least one bearing system and includes one or more airfoils. 
     
     
         18 . The gas turbine engine of  claim 15 , wherein the low corrected fan tip speed is less than about 1150 ft/sec. 
     
     
         19 . The gas turbine engine of  claim 15 , wherein the low pressure turbine includes at least one rotor constrained by a first stress level, at least one of the plurality of fan blades of the fan constrained by a second stress level and having a fan tip speed boundary condition, and the star gear ratio is configured such that in operation the fan blade does not exceed the fan tip speed boundary condition or the second stress level, and the low pressure turbine rotor does not exceed the first stress level. 
     
     
         20 . A gas turbine engine comprising:
 an engine centerline longitudinal axis;   a fan section including a fan with a plurality of fan blades and rotatable about the engine centerline longitudinal axis;   a low corrected fan tip speed less than about 1150 ft/sec, wherein the low corrected fan tip speed is an actual fan tip speed at an ambient temperature divided by [(Tram° R)/(518.7° R)] 0.5 , where T represents the ambient temperature in degrees Rankine;   a bypass ratio of greater than about 11.0;   a speed reduction device comprising a gear system;   a plurality of bearing systems;   a first spool comprising a first turbine and a first shaft; and   a second spool comprising a second turbine and a second shaft;   wherein the first shaft and second shaft are concentric and mounted via at least one of the plurality of bearing systems for rotation about the engine centerline longitudinal axis, and the first shaft is in communication with the fan through the speed reduction device; and   wherein the first turbine includes at least three stages and no more than four stages and a pressure ratio greater than about 5:1, the first turbine includes an inlet having an inlet pressure, and an outlet that is prior to any exhaust nozzle and having an outlet pressure, and the pressure ratio of the first turbine is a ratio of the inlet pressure to the outlet pressure.   
     
     
         21 . The gas turbine engine of  claim 20 , wherein the second turbine includes two stages. 
     
     
         22 . The gas turbine engine of  claim 21 , wherein the first turbine includes three stages. 
     
     
         23 . The gas turbine engine of  claim 21 , wherein the gear system has a gear ratio of less than or equal to 4.1. 
     
     
         24 . The gas turbine engine of  claim 23 , wherein the gear reduction ratio is at least 2.6. 
     
     
         25 . The gas turbine engine of  claim 24 , wherein the gear system is a star gear system with a ring gear, and a sun gear, wherein the gear ratio is determined by measuring a diameter of the ring gear and dividing that diameter by the diameter of the sun gear. 
     
     
         26 . The gas turbine engine of  claim 24 , wherein the first turbine includes four stages. 
     
     
         27 . A gas turbine engine comprising:
 an engine centerline longitudinal axis;   a fan section including a fan with a plurality of fan blades and rotatable about the engine centerline longitudinal axis;   a low corrected fan tip speed less than about 1400 ft/sec, wherein the low corrected fan tip speed is an actual fan tip speed at an ambient temperature divided by [(Tram° R)/(518.7° R)] 0.5 , where T represents the ambient temperature in degrees Rankine;   a bypass ratio greater than about 11;   a speed reduction device comprising a gear system with a gear ratio;   a plurality of bearing systems;   a low speed spool comprising a low pressure turbine and a first shaft; and   a high speed spool comprising a high pressure turbine and a second shaft;   wherein the first shaft and second shaft are concentric and mounted via at least one of the plurality of bearing systems for rotation about the engine centerline longitudinal axis, and the first shaft is in communication with the fan through the speed reduction device; and   wherein the low pressure turbine includes four stages.   
     
     
         28 . The gas turbine engine of  claim 27 , wherein the high pressure turbine includes two stages. 
     
     
         29 . The gas turbine engine of  claim 27 , wherein the gear reduction ratio is at least 2.6 and less than or equal to 4.1. 
     
     
         30 . The gas turbine engine of  claim 29 , wherein the low pressure turbine has a pressure ratio greater than about 5:1, the low pressure turbine includes an inlet having an inlet pressure, and an outlet that is prior to any exhaust nozzle and having an outlet pressure, and the pressure ratio of the low pressure turbine is a ratio of the inlet pressure to the outlet pressure.