FLEXIBLE SUPPORT STRUCTURE FOR A GEARED ARCHITECTURE GAS TURBINE ENGINE
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 that supports the fan shaft and a gear system connected to the fan shaft. The gear system includes a ring gear defining a ring gear transverse stiffness, a gear mesh defining a gear mesh transverse stiffness, and a reduction ratio greater than 2.3. The ring gear transverse stiffness is less than 20% 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;
a gear system connected to said fan shaft, said gear system includes a ring gear defining a ring gear transverse stiffness, a gear mesh defining a gear mesh transverse stiffness, and a reduction ratio greater than 2.3; and
wherein said ring gear transverse stiffness is less than 20% of 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 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 , further comprising a two stage high pressure turbine.
5 . The gas turbine engine of claim 2 , 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.
6 . The gas turbine engine of claim 2 , further comprising a gear system input to said gear system defining a gear system input transverse stiffness and said fan shaft support defining a fan shaft support transverse stiffness, wherein said gear system input transverse stiffness is less than 20% of 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 said gear mesh defines a gear mesh lateral stiffness and said gear system input defines a gear system input lateral stiffness and said gear system input lateral stiffness is less than 5% of said gear mesh lateral stiffness.
9 . The gas turbine engine of claim 8 , wherein said fan shaft support defines a fan shaft support lateral stiffness and said gear system input lateral stiffness is less than 11% of said fan shaft support lateral 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 said gear system input transverse stiffness is less than 11% of said fan shaft support transverse stiffness.
13 . The gas turbine engine of claim 11 , 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.
14 . The gas turbine engine of claim 13 , wherein said gear system input transverse stiffness is less than 11% of said fan shaft support transverse stiffness, and said gear system input both transfers torque and facilitates segregation of vibrations.
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 a ring gear transverse stiffness, a gear mesh defining a gear mesh transverse stiffness, and a reduction ratio greater than 2.3;
wherein said ring gear transverse stiffness is less than 20% of 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 said ring gear transverse stiffness is less than 20% of said gear mesh transverse stiffness and said ring gear lateral stiffness is less than 12% of said gear mesh lateral 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 blades alone.
22 . The gas turbine engine of claim 21 , wherein said ring gear transverse stiffness is less than 12% of said gear mesh transverse stiffness.
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 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.
25 . The gas turbine engine of claim 24 , wherein said ring gear transverse stiffness is less than 12% of said gear mesh transverse stiffness.
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.