Geared architecture for high speed and small volume fan drive turbine

Abstract

A gas turbine engine according to an example of the present disclosure includes, among other things, a fan shaft driving a fan having fan blades, a fan shaft support said that supports fan shaft and a gear system connected to the fan shaft. The gear system includes a ring gear defining a ring gear lateral stiffness and a ring gear transverse stiffness, a gear mesh defining a gear mesh lateral stiffness and a gear mesh transverse stiffness, and a reduction ratio greater than 2.3. At least one of the ring gear lateral stiffness and the ring gear transverse stiffness is less than 12% of a respective one of the gear mesh lateral stiffness and the gear mesh transverse stiffness.

What is claimed is: 
     
         1 . A gas turbine engine, comprising:
 a fan shaft driving a fan having fan blades;   a fan shaft support that supports said fan shaft;   a gear system connected to said fan shaft, said gear system includes a ring gear defining a ring gear lateral stiffness and a ring gear transverse stiffness, a gear mesh defining a gear mesh lateral stiffness and a gear mesh transverse stiffness, and a reduction ratio greater than 2.3; and   wherein at least one of said ring gear lateral stiffness and said ring gear transverse stiffness is less than 12% of a respective one of said gear mesh lateral stiffness and said gear mesh transverse stiffness.   
     
     
         2 . The gas turbine engine of  claim 1 , further comprising a bypass ratio greater than ten (10). 
     
     
         3 . The gas turbine engine of  claim 2 , further comprising a two stage high pressure turbine, and a low corrected fan tip speed less than about 1150 ft/second, wherein said 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 said ambient temperature in degrees Rankine. 
     
     
         4 . The gas turbine engine of  claim 3 , further comprising a low pressure turbine, and a three stage low pressure compressor. 
     
     
         5 . The gas turbine engine of  claim 2 , further comprising a low pressure turbine with an inlet, an outlet, and a low pressure turbine pressure ratio greater than 5:1, wherein said low pressure turbine pressure ratio is a ratio of a pressure measured prior to said inlet as related to a pressure at said outlet prior to any exhaust nozzle. 
     
     
         6 . The gas turbine engine of  claim 2 , further comprising an input coupling to said gear system defining an input coupling lateral stiffness and an input coupling transverse stiffness and said fan shaft support defining a fan shaft support lateral stiffness and a fan shaft support transverse stiffness, wherein at least one of said input coupling lateral stiffness and said input coupling transverse stiffness is less than 11% of a respective one of said fan shaft support lateral stiffness and said fan shaft support transverse stiffness. 
     
     
         7 . The gas turbine engine of  claim 6 , further comprising a mid-turbine frame including at least one airfoil extending into a flow path, and wherein said fan shaft support is a k-frame bearing support. 
     
     
         8 . The gas turbine engine of  claim 6 , wherein at least one of said input coupling lateral stiffness and said input coupling transverse stiffness is less than 5% of a respective one of said gear mesh lateral stiffness and said gear mesh transverse stiffness. 
     
     
         9 . The gas turbine engine of  claim 8 , wherein both said input coupling lateral stiffness and said input coupling transverse stiffness are less than 11% of a respective one of a fan shaft support lateral stiffness and a fan shaft support transverse stiffness. 
     
     
         10 . The gas turbine engine of  claim 6 , further comprising a two stage high pressure turbine and a low fan pressure ratio of less than 1.45 and said low fan pressure ratio is measured across the fan blades alone. 
     
     
         11 . The gas turbine engine of  claim 10 , further comprising a low pressure turbine with an inlet, an outlet, and a low pressure turbine pressure ratio greater than 5:1, wherein said low pressure turbine pressure ratio is a ratio of a pressure measured prior to said inlet as related to a pressure at said outlet prior to any exhaust nozzle. 
     
     
         12 . The gas turbine engine of  claim 11 , wherein both said input coupling lateral stiffness and said input coupling transverse stiffness are less than 11% of a respective one of said fan shaft support lateral stiffness and said fan shaft support lateral stiffness. 
     
     
         13 . The gas turbine engine of  claim 11 , wherein said input coupling both transfers torque and facilitates segregation of vibrations. 
     
     
         14 . The gas turbine engine of  claim 13 , further comprising a low corrected fan tip speed less than about 1150 ft/second, wherein said 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 said ambient temperature in degrees Rankine. 
     
     
         15 . A gas turbine engine, comprising:
 a fan shaft driving a fan having fan blades;   a fan shaft support that supports said fan shaft;   a gear system connected to said fan shaft, said gear system includes a ring gear defining ring gear lateral stiffness and a ring gear transverse stiffness, a gear mesh defining a gear mesh lateral stiffness and a gear mesh transverse stiffness, and a reduction ratio greater than 2.3; and   wherein at least one of said ring gear lateral stiffness and said ring gear transverse stiffness is less than 12% of a respective one of said gear mesh lateral stiffness and said gear mesh transverse stiffness; and   a low fan pressure ratio of less than 1.45 and said low fan pressure ratio is measured across the fan blades alone.   
     
     
         16 . The gas turbine engine of  claim 15 , further comprising a two stage high pressure turbine and a low pressure turbine with an inlet, an outlet, and a low pressure turbine pressure ratio greater than 5:1, wherein said low pressure turbine pressure ratio is a ratio of a pressure measured prior to said inlet as related to a pressure at said outlet prior to any exhaust nozzle. 
     
     
         17 . The gas turbine engine of  claim 16 , further comprising a low corrected fan tip speed less than about 1150 ft/second, wherein said 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 said ambient temperature in degrees Rankine. 
     
     
         18 . The gas turbine engine of  claim 17 , further comprising a mid-turbine frame including at least one airfoil extending into a flow path. 
     
     
         19 . The gas turbine engine of  claim 15 , further comprising a low corrected fan tip speed less than about 1150 ft/second, wherein said 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 said ambient temperature in degrees Rankine. 
     
     
         20 . A gas turbine engine, comprising:
 a fan shaft driving a fan having fan blades;   a fan shaft support that supports said fan shaft;   a gear system connected to said fan shaft, said gear system includes a ring gear defining a ring gear transverse stiffness and a ring gear lateral stiffness, a gear mesh defining a gear mesh transverse stiffness and a gear mesh lateral stiffness, and a reduction ratio greater than 2.3; and   wherein both of said ring gear lateral stiffness and said ring gear transverse stiffness are less than 12% of a respective one of said gear mesh lateral stiffness and said gear mesh transverse stiffness.   
     
     
         21 . The gas turbine engine of  claim 20 , further comprising a low fan pressure ratio of less than 1.45 and said low fan pressure ratio is measured across the fan blade alones. 
     
     
         22 . The gas turbine engine of  claim 21 , further comprising at least one bearing system, a high speed spool including an outer shaft, and a low speed spool including an inner shaft, wherein said gear system includes a planet carrier and said fan shaft is mounted to said planet carrier and said inner shaft and outer shaft are concentric and rotate via said at least one bearing system about a longitudinal axis of said engine. 
     
     
         23 . The gas turbine engine of  claim 21 , further comprising a low pressure turbine with an inlet, an outlet, and a low pressure turbine pressure ratio greater than 5:1, wherein said low pressure turbine pressure ratio is a ratio of a pressure measured prior to said inlet as related to a pressure at said outlet prior to any exhaust nozzle. 
     
     
         24 . The gas turbine engine of  claim 23 , further comprising at least one bearing system, a high speed spool, a low speed spool, and a low corrected fan tip speed less than about 1150 ft/second, wherein said 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 said ambient temperature in degrees Rankine, and said high speed spool includes an outer shaft and said low speed spool includes an inner shaft, and said inner shaft and outer shaft are concentric and rotate in coordination with said at least one bearing system about a longitudinal axis of said engine. 
     
     
         25 . The gas turbine engine of  claim 24 , further comprising a two stage high pressure turbine, wherein said gear system includes a planet carrier and said fan shaft is mounted to said planet carrier. 
     
     
         26 . The gas turbine engine of  claim 25 , further comprising a mid-turbine frame including at least one airfoil extending into a flow path. 
     
     
         27 . The gas turbine engine of  claim 21 , further comprising a low corrected fan tip speed less than about 1150 ft/second, wherein said 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 said ambient temperature in degrees Rankine. 
     
     
         28 . The gas turbine engine of  claim 20 , further comprising a high speed spool including a two stage high pressure turbine, and a low speed spool including a low pressure turbine with an inlet, an outlet, and a low pressure turbine pressure ratio greater than 5:1, wherein said low pressure turbine pressure ratio is a ratio of a pressure measured prior to said inlet as related to a pressure at said outlet prior to any exhaust nozzle. 
     
     
         29 . The gas turbine engine of  claim 28 , wherein said gear system includes a planet carrier and said fan shaft is mounted to said planet carrier, and further comprising a low corrected fan tip speed less than about 1150 ft/second, wherein said 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 said ambient temperature in degrees Rankine. 
     
     
         30 . The gas turbine engine of  claim 29 , further comprising at least one bearing system, and a three stage low pressure compressor, wherein said high speed spool includes an outer shaft and said low speed spool includes an inner shaft, and said inner shaft and outer shaft are concentric and rotate in coordination with said at least one bearing system about a longitudinal axis of said engine.