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 an engine centerline longitudinal axis. A fan section includes a fan with a plurality of fan blades. The fan has a low corrected fan tip speed less than 1400 ft/sec. A bypass ratio is greater than 13 and less than 20. A fan pressure ratio less than 1.38 at cruise conditions of 0.8 Mach and about 35,000 feet. A speed reduction device comprises a gear system with a gear ratio of at least 2.6 and less than or equal to 4.1. A low pressure turbine is in communication with a first shaft. A high pressure turbine is in communication a second shaft. The first shaft is in communication with the fan through the speed reduction device. The low pressure turbine includes at least three stages and no more than four stages. The high pressure turbine includes two 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 wherein the fan section drives air along a bypass flow path in a bypass duct and the fan rotates about the engine centerline longitudinal axis;   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 T represents the ambient temperature in degrees Rankine;   a fan pressure ratio less than 1.38 at cruise conditions of 0.8 Mach and about 35,000 feet, wherein the fan pressure ratio is measured across a fan blade alone;   a speed reduction device comprising an epicyclic gear system having five intermediate gears and a gear ratio of at least 2.6 and less than or equal to 4.1;   a plurality of bearing systems;   a low pressure turbine in communication with a first shaft;   a high pressure turbine in communication a second shaft; and   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 one or more airfoils that extend in a flow path; and   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;   wherein the high pressure turbine includes two 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 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.   
     
     
         2 . The gas turbine engine of  claim 1 , wherein the low pressure turbine includes four stages. 
     
     
         3 . The gas turbine engine of  claim 2 , 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. 
     
     
         4 . The gas turbine engine of  claim 2 , a bypass ratio greater than 11.0 and less than 22.0 at cruise conditions of 0.8 Mach and about 35,000 feet. 
     
     
         5 . The gas turbine engine of  claim 2 , a bypass ratio greater than 13 and less than 20 at cruise conditions of 0.8 Mach and about 35,000 feet. 
     
     
         6 . A gas turbine engine comprising:
 an engine centerline longitudinal axis;   a fan section including a fan with a plurality of fan blades wherein the fan section drives air along a bypass flow path in a bypass duct and the fan rotates about the engine centerline longitudinal axis;   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 T represents the ambient temperature in degrees Rankine;   a fan pressure ratio less than 1.38 at cruise conditions of 0.8 Mach and about 35,000 feet, wherein the fan pressure ratio is measured across a fan blade alone;   a speed reduction device comprising an epicyclic gear system having five intermediate gears;   a plurality of bearing systems;   a low pressure turbine in communication with a first shaft;   a high pressure turbine in communication with a second shaft; and   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 one or more airfoils that extend in a flow path; and   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 high pressure turbine includes two 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 system 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.   
     
     
         7 . The gas turbine engine of  claim 6 , wherein the low pressure turbine includes five stages. 
     
     
         8 . The gas turbine engine of  claim 6 , further comprising a low pressure compressor including three stages, wherein the low pressure turbine drives the low pressure compressor. 
     
     
         9 . The gas turbine engine of  claim 6 , wherein the low pressure turbine includes at least three stages and no more than four stages. 
     
     
         10 . The gas turbine engine of  claim 6 , wherein the speed reduction device includes a gear ratio less than or equal to 4.1. 
     
     
         11 . The gas turbine engine of  claim 10 , wherein the gear ratio is greater than or equal to 2.6. 
     
     
         12 . The gas turbine engine of  claim 9 , wherein the low pressure turbine includes four stages. 
     
     
         13 . The gas turbine engine of  claim 12 , a bypass ratio greater than 11.0 and less than 22.0. 
     
     
         14 . The gas turbine engine of  claim 13 , wherein the bypass ratio is at cruise conditions of 0.8 Mach and about 35,000 feet. 
     
     
         15 . The gas turbine engine of  claim 12 , further comprising a low pressure compressor including three stages, wherein the low pressure turbine drives the low pressure compressor. 
     
     
         16 . The gas turbine engine of  claim 15 , further comprising a high pressure compressor including eight stages, wherein the high pressure turbine drives the high pressure compressor. 
     
     
         17 . The gas turbine engine of  claim 15 , 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. 
     
     
         18 . The gas turbine engine of  claim 17 , wherein the gear system includes a carrier supporting the intermediate gears and fixed from rotation relative to a static structure of the gas turbine engine. 
     
     
         19 . The gas turbine engine of  claim 6 , a bypass ratio greater than 13 and less than 20 at cruise conditions of 0.8 Mach and about 35,000 feet. 
     
     
         20 . A gas turbine engine comprising:
 an engine centerline longitudinal axis;   a fan section including a fan with a plurality of fan blades wherein the fan section drives air along a bypass flow path in a bypass duct and the fan rotates about the engine centerline longitudinal axis;   a low corrected fan tip speed less than 1400 ft/sec, and 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 11.0 and less than 22 at cruise conditions of 0.8 Mach and about 35,000 feet;   a fan pressure ratio less than 1.38 and greater than 1.25 at cruise conditions of 0.8 Mach and about 35,000 feet, wherein the fan pressure ratio is measured across a fan blade alone;   a speed reduction device comprising an epicyclic gear system having five intermediate gears;   a plurality of bearing systems;   a low pressure turbine in communication with a first shaft and includes a pressure ratio greater than about 5:1, the low pressure turbine includes an inlet having an inlet pressure, and an outlet having an outlet pressure, and the pressure ratio of the low pressure turbine is a ratio of the inlet pressure to the outlet pressure;   a high pressure turbine in communication with a second shaft; and   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 one or more airfoils that extend in a flow path; and   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   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 system 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.   
     
     
         21 . The gas turbine engine of  claim 20 , wherein the high pressure turbine includes two stages. 
     
     
         22 . The gas turbine engine of  claim 21 , wherein the low pressure 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 low pressure 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 wherein the fan section drives air along a bypass flow path in a bypass duct and the fan rotates about the engine centerline longitudinal axis;   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 T represents the ambient temperature in degrees Rankine;   a fan pressure ratio less than 1.38 and greater than 1.25 at cruise conditions of 0.8 Mach and about 35,000 feet, wherein the fan pressure ratio is measured across a fan blade alone;   a speed reduction device comprising an epicyclic gear system having five intermediate gears and a gear ratio of at least 2.6 and less than or equal to 4.1;   a plurality of bearing systems;   a low pressure turbine in communication with a first shaft;   a high pressure turbine in communication with a second shaft;   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 one or more airfoils that extend in a flow path; and   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 four 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 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.   
     
     
         28 . The gas turbine engine of  claim 27 , wherein the high pressure turbine includes two stages. 
     
     
         29 . The gas turbine engine of  claim 27 , a bypass ratio greater than 11 and less than 22 at cruise conditions of 0.8 Mach and about 35,000 feet. 
     
     
         30 . The gas turbine engine of  claim 29 , wherein the low pressure turbine has a pressure ratio greater than 5:1, the low pressure turbine includes an inlet having an inlet pressure, and an outlet having an outlet pressure, and the pressure ratio of the low pressure turbine is a ratio of the inlet pressure to the outlet pressure.