Flexible support structure for a geared architecture gas turbine engine
A gas turbine engine includes a fan shaft that drives a fan that has fan blades. A fan shaft support that supports the fan shaft defines a fan shaft support transverse stiffness. A gear system is connected to the fan shaft and includes a gear mesh that defines a gear mesh transverse stiffness. A flexible support supports the gear system and defines a flexible support transverse stiffness. The flexible support transverse stiffness is less than 11% of the fan shaft support transverse stiffness and less than 8% of the gear mesh transverse stiffness.
1. A gas turbine engine, comprising:
a fan shaft driving a fan having fan blades;
a fan shaft support that supports said fan shaft defining a fan shaft support transverse stiffness;
a gear system connected to said fan shaft, wherein said gear system includes a gear mesh defining a gear mesh transverse stiffness;
a flexible support supporting said gear system defining a flexible support transverse stiffness; and
wherein said flexible support transverse stiffness is less than 11% of said fan shaft support transverse stiffness and less than 8% of said gear mesh transverse stiffness.
2. The gas turbine engine of claim 1 , wherein said gear mesh defines a gear mesh lateral stiffness and said gear system includes a ring gear defining a ring gear lateral stiffness and said ring gear lateral stiffness is less than 12% of said gear mesh lateral stiffness.
3. The gas turbine engine of claim 2 , wherein said flexible support defines a flexible support lateral stiffness and said fan shaft support defines a fan shaft support lateral stiffness and said flexible support lateral stiffness is less than 11% of said fan shaft support lateral stiffness.
4. The gas turbine engine of claim 3 , wherein said gear system is an epicyclic gear system, said gear mesh defines a gear mesh lateral stiffness and said flexible support lateral stiffness is less than 8% of said gear mesh lateral 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 , further comprising a low fan pressure ratio of less than 1.45 and said low fan pressure ratio measured across the fan blades alone.
7. The gas turbine engine of claim 6 , further comprising a bypass ratio greater than ten (10).
8. The gas turbine engine of claim 7 , 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.
9. The gas turbine engine of claim 8 , wherein said gear system includes a reduction ratio greater than 2.3.
10. The gas turbine engine of claim 9 , further comprising a low corrected fan tip speed less than 1150 ft/second, wherein said low corrected fan tip speed is an actual fan tip speed at an ambient temperature divided by [(T ram °R)/(518.7 °R)] 0.5 , where T represents said ambient temperature in degrees Rankine.
11. The gas turbine engine of claim 5 , wherein said gear system includes a reduction ratio greater than 2.3.
12. The gas turbine engine of claim 11 , further comprising a low fan pressure ratio of less than 1.45 and said low fan pressure ratio measured across the fan blades alone.
13. The gas turbine engine of claim 12 , further comprising a low corrected fan tip speed less than 1150 ft/second, wherein said low corrected fan tip speed is an actual fan tip speed at an ambient temperature divided by [(T ram °R)/(518.7 °R)] 0.5 , where T represents said ambient temperature in degrees Rankine.
14. The gas turbine engine of claim 13 , further comprising a bypass ratio greater than ten (10).
15. The gas turbine engine of claim 14 , 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.
16. 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 gear mesh defining a gear mesh transverse stiffness;
a flexible support supporting said gear system defining a flexible support transverse stiffness; and
wherein said flexible support transverse stiffness is less than 8% of said gear mesh transverse stiffness.
17. The gas turbine engine of claim 16 , wherein said gear system includes a ring gear defining a ring gear transverse stiffness that is less than 12% of said gear mesh transverse stiffness.
18. The gas turbine engine of claim 17 , wherein said fan shaft support defines a fan shaft support transverse stiffness and said flexible support transverse stiffness is less than 11% of said fan shaft support transverse stiffness.
19. The gas turbine engine of claim 18 , wherein said fan shaft support defines a fan shaft support lateral stiffness and said flexible support defines a flexible support lateral stiffness that is less than 11% of said fan shaft support lateral stiffness.
20. The gas turbine engine of claim 19 , wherein said gear mesh defines a gear mesh lateral stiffness and said flexible support lateral stiffness is less than 8% of said gear mesh lateral stiffness.
21. The gas turbine engine of claim 20 , 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.
22. The gas turbine engine of claim 21 , wherein said ring gear defines a ring gear lateral stiffness that is less than 12% of said gear mesh lateral stiffness.
23. The gas turbine engine of claim 21 , further comprising a low fan pressure ratio of less than 1.45 and said low fan pressure ratio measured across the fan blades alone.
24. The gas turbine engine of claim 23 , further comprising a bypass ratio greater than ten (10).
25. The gas turbine engine of claim 24 , wherein said gear system is an epicyclic gear system, and said gear system includes a reduction ratio greater than 2.3.
26. The gas turbine engine of claim 25 , further comprising a low corrected fan tip speed less than 1150 ft/second, wherein said low corrected fan tip speed is an actual fan tip speed at an ambient temperature divided by [(T ram °R)/(518.7 °R)] 0.5 , where T represents said ambient temperature in degrees Rankine.
27. The gas turbine engine of claim 20 , further comprising a bypass ratio greater than ten (10).
28. The gas turbine engine of claim 27 , wherein said ring gear defines a ring gear lateral stiffness that is less than 12% of said gear mesh lateral stiffness.
29. The gas turbine engine of claim 28 , further comprising a low fan pressure ratio of less than 1.45, said low fan pressure ratio measured across the fan blades alone.
30. The gas turbine engine of claim 29 , further comprising a low corrected fan tip speed less than 1150 ft/second, wherein said low corrected fan tip speed is an actual fan tip speed at an ambient temperature divided by [(T ram °R)/(518.7 °R)] 0.5 , where T represents said ambient temperature in degrees Rankine.