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

Abstract

In one exemplary embodiment, a gas turbine engine includes a fan section having a fan with a low corrected fan tip speed less than 1400 ft/sec. A bypass ratio is greater than 11.0 and less than 22.0. A fan pressure ratio is less than 1.48. A speed reduction device includes a gear system with a gear ratio of at least 2.6 and less than or equal to 4.1. A low pressure turbine including three stages and a high pressure turbine including two stages. The low pressure turbine includes at least one rotor constrained by a first stress level. At least one of a plurality of fan blades is constrained by a second stress level and has a fan tip speed boundary condition. The gear ratio is configured such that the at least one fan blade does not exceed the fan tip speed boundary condition or the second stress level.

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  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 Tram represents the ambient temperature in degrees Rankine;   a bypass ratio greater than 11.0 and less than 22.0;   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 low pressure turbine in communication with a first shaft;   a high pressure turbine in communication with a second shaft;   wherein the first shaft and second shaft are concentric and mounted via at least one of a 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;   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 at least one of the plurality of fan blades does not exceed the fan tip speed boundary condition or the second stress level, and the at least one rotor does not exceed the first stress level; 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 at least four stages and no more than eight 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 is a four-stage low pressure turbine. 
     
     
         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 11.0 and less than 22.0;   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 low pressure turbine in communication with a first shaft; and   a high pressure turbine in communication with a second shaft;   wherein the first shaft and second shaft are concentric and mounted via at least one of a 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 last three stages;   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 at least one of the plurality of fan blades does not exceed the fan tip speed boundary condition or the second stress level, and the at least one rotor does not exceed the first stress level; and   wherein the high pressure turbine includes two stages.   
     
     
         8 . The gas turbine engine of  claim 7 , wherein the low pressure turbine is a five-stage low pressure turbine. 
     
     
         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 is a four-stage low pressure turbine. 
     
     
         12 . The gas turbine engine of  claim 11 , further comprising a low corrected fan tip speed that does not exceed 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 Tram 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 of the plurality of bearing systems and includes one or more airfoils. 
     
     
         18 . The gas turbine engine of  claim 15 , wherein the low corrected fan tip speed is less than 1150 ft/sec. 
     
     
         19 . 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 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 Tram represents the ambient temperature in degrees Rankine;   a bypass ratio of greater than 11.0 and less than 22.0;   a speed reduction device comprising a gear system with a gear reduction ratio;   a first turbine in communication with a first shaft; and   a second turbine in communication with a second shaft;   wherein the first shaft and second shaft are concentric and mounted via at least one of a 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 first turbine includes at least three stages and no more than four stages; and   wherein the first 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 reduction ratio is configured such that in operation the at least one of the plurality of fan blades does not exceed the fan tip speed boundary condition or the second stress level, and the at least one rotor does not exceed the first stress level.   
     
     
         20 . The gas turbine engine of  claim 19 , wherein the second turbine is a two-stage second turbine. 
     
     
         21 . The gas turbine engine of  claim 20 , wherein the first turbine is a three-stage first turbine. 
     
     
         22 . The gas turbine engine of  claim 20 , 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 and the gear reduction ratio is less than or equal to 4.1. 
     
     
         23 . The gas turbine engine of  claim 22 , wherein the gear reduction ratio is at least 2.6. 
     
     
         24 . The gas turbine engine of  claim 23 , wherein the gear system is a star gear system with a ring gear, and a sun gear, wherein the gear reduction ratio is determined by measuring a diameter of the ring gear and dividing that diameter by the diameter of the sun gear. 
     
     
         25 . The gas turbine engine of  claim 23 , wherein the first turbine is a four-stage first turbine. 
     
     
         26 . 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 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 Tram represents the ambient temperature in degrees Rankine;   a bypass ratio greater than 11 and less than 22;   a speed reduction device comprising a gear system with a gear reduction ratio;   a low pressure turbine in communication with a first shaft;   a high pressure turbine in communication with a second shaft;   wherein the first shaft and second shaft are concentric and mounted via at least one of a 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 8 stages; and   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 reduction 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 at least one rotor does not exceed the first stress level.   
     
     
         27 . The gas turbine engine of  claim 26 , wherein the high pressure turbine includes two stages. 
     
     
         28 . The gas turbine engine of  claim 26 , wherein the gear reduction ratio is at least 2.6 and less than or equal to 4.1. 
     
     
         29 . The gas turbine engine of  claim 28 , 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.