Flexible support structure for a geared architecture gas turbine engine

Abstract

A gas turbine engine includes a fan that has fan blades. A fan shaft is drivingly connected to the fan. A gear system is connected to the fan shaft and driven through an input. A gear system flex mount arrangement accommodates misalignment of the fan shaft and the input during operation. The gear system flex mount arrangement includes a gear mesh that defines a gear mesh lateral stiffness and a flexible support that defines a flexible support lateral stiffness that is less than 8% of the gear mesh lateral stiffness.

What is claimed is: 
     
         1 . A gas turbine engine, comprising:
 a fan having fan blades;   a fan shaft drivingly connected to said fan;   a gear system connected to said fan shaft and driven through an input; and   a gear system flex mount arrangement, wherein said gear system flex mount arrangement accommodates misalignment of said fan shaft and said input during operation and said gear system flex mount arrangement includes a gear mesh defining a gear mesh lateral stiffness and a flexible support defining a flexible support lateral stiffness that is less than 8% of said gear mesh lateral stiffness.   
     
     
         2 . The gas turbine engine of  claim 1 , wherein said gear system flex mount arrangement further includes a ring gear defining a ring gear lateral stiffness that is less than 12% of said gear mesh lateral stiffness. 
     
     
         3 . The gas turbine engine of  claim 2 , wherein said gear system has a gear reduction ratio of greater than 2.3, and further comprising a bypass ratio greater than ten (10), a fan pressure ratio of less than 1.45 measured across said fan blades alone, 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. 
     
     
         4 . The gas turbine engine of  claim 3 , wherein said gear mesh defines a gear mesh transverse stiffness and said flexible support defines a flexible support transverse stiffness that is less than 8% of said gear mesh transverse stiffness. 
     
     
         5 . The gas turbine engine of  claim 4 , wherein said ring gear defines a ring gear transverse stiffness that is less than 12% of said gear mesh transverse stiffness. 
     
     
         6 . The gas turbine engine of  claim 5 , wherein said gear system flex mount arrangement further includes a frame supporting said fan shaft and defining a frame lateral stiffness, wherein said flexible support lateral stiffness is less than 11% of said frame lateral stiffness. 
     
     
         7 . The gas turbine engine of  claim 6 , further comprising a two stage high pressure turbine. 
     
     
         8 . The gas turbine engine of  claim 1 , wherein said gear system flex mount arrangement further includes a frame supporting said fan shaft and defining a frame lateral stiffness, wherein said flexible support lateral stiffness is less than 11% of said frame lateral stiffness. 
     
     
         9 . The gas turbine engine of  claim 8 , wherein said flexible support defines a flexible support transverse stiffness and said gear mesh defines a gear mesh transverse stiffness and said flexible support transverse stiffness is less than 8% of said gear mesh transverse stiffness. 
     
     
         10 . The gas turbine engine of  claim 9 , wherein said frame defines a frame transverse stiffness and said flexible support transverse stiffness is less than 11% of gear frame transverse stiffness. 
     
     
         11 . The gas turbine engine of  claim 10 , wherein said gear system flex mount arrangement further includes a ring gear defining a ring gear transverse stiffness that is less than 20% of said gear mesh transverse stiffness. 
     
     
         12 . The gas turbine engine of  claim 11 , wherein said input defines an input transverse stiffness that is less than 20% of said frame transverse stiffness. 
     
     
         13 . The gas turbine engine of  claim 12 , wherein said ring gear defines a ring gear lateral stiffness that is less than 12% of said gear mesh lateral stiffness. 
     
     
         14 . The gas turbine engine of  claim 13 , wherein said input defines an input lateral stiffness that is less than 11% of said frame lateral stiffness. 
     
     
         15 . The gas turbine engine of  claim 14 , further comprising a two stage high pressure turbine. 
     
     
         16 . The gas turbine engine of  claim 15 , wherein said gear system has a gear reduction ratio of greater than 2.3, and further comprising a bypass ratio greater than ten (10), a fan pressure ratio of less than 1.45 measured across said fan blades alone, 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 . A gas turbine engine, comprising:
 a fan shaft having fan blades;   a fan shaft drivingly connected to said fan;   a gear system connected to said fan shaft and driven through an input defining an input lateral stiffness; and   a gear system flex mount arrangement, wherein said gear system flex mount arrangement accommodates misalignment of said fan shaft and said input during operation and includes a frame for supporting said fan shaft defining a frame lateral stiffness and said input lateral stiffness is less than 11% of said frame lateral stiffness.   
     
     
         18 . The gas turbine engine of  claim 17 , wherein said gear system flex mount arrangement further includes a gear mesh defining a gear mesh lateral stiffness and said input lateral stiffness is less than 5% of said gear mesh lateral stiffness. 
     
     
         19 . The gas turbine engine of  claim 18 , wherein said gear system flex mount arrangement further includes a ring gear defining a ring gear lateral stiffness that is less than 12% of said gear mesh lateral stiffness. 
     
     
         20 . The gas turbine engine of  claim 19 , wherein said gear system has a gear reduction ratio of greater than 2.3, and further comprising a bypass ratio greater than ten (10), a fan pressure ratio of less than 1.45 measured across said fan blades alone, 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. 
     
     
         21 . The gas turbine engine of  claim 20 , wherein said gear system flex mount arrangement further includes a flexible support defining a flexible support transverse stiffness, said frame defines a frame transverse stiffness, and said flexible support transverse stiffness is less than 11% of said frame transverse stiffness. 
     
     
         22 . The gas turbine engine of  claim 21 , further comprising a two stage high pressure turbine. 
     
     
         23 . The gas turbine engine of  claim 18 , wherein said gear system flex mount arrangement further includes a flexible support defining a flexible support lateral stiffness that is less than 11% of said frame lateral stiffness. 
     
     
         24 . The gas turbine engine of  claim 17 , wherein said gear system flex mount arrangement further includes a ring gear defining a ring gear lateral stiffness and a gear mesh defining a gear mesh lateral stiffness and said ring gear lateral stiffness is less than 12% of said gear mesh lateral stiffness. 
     
     
         25 . The gas turbine engine of  claim 24 , wherein said input lateral stiffness is less than 5% of said gear mesh lateral stiffness. 
     
     
         26 . The gas turbine engine of  claim 25 , wherein said ring gear define a ring gear transverse stiffness and said gear mesh defines a gear mesh transverse stiffness and said ring gear transverse stiffness is less than 12% of said gear mesh transverse stiffness. 
     
     
         27 . The gas turbine engine of  claim 17 , wherein said gear system flex mount arrangement includes a flexible support defining a flexible support lateral stiffness that is less than 11% of said frame lateral stiffness. 
     
     
         28 . The gas turbine engine of  claim 27 , wherein said input defines an input transverse stiffness, said frame defines a frame transverse stiffness, said gear system includes a gear mesh defining a gear mesh transverse stiffness, and said input transverse stiffness is less than 11% of said frame transverse stiffness and less than 5% of said gear mesh transverse stiffness. 
     
     
         29 . The gas turbine engine of  claim 28 , wherein said gear system flex mount arrangement further includes a ring gear defining a ring gear lateral stiffness, said gear mesh defines a gear mesh lateral stiffness, and said ring gear lateral stiffness is less than 12% of said gear mesh lateral stiffness. 
     
     
         30 . The gas turbine engine of  claim 27 , wherein said gear system has a gear reduction ratio of greater than 2.3, and further comprising a bypass ratio greater than ten (10), a fan pressure ratio of less than 1.45 measured across said fan blades alone, 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.