TURBINE SECTION OF HIGH BYPASS TURBOFAN
A turbofan engine has an engine case and a gaspath through the engine case. A fan has a circumferential array of fan blades. The engine further has a compressor, a combustor, a gas generating turbine, and a low pressure turbine section. A speed reduction mechanism couples the low pressure turbine section to the fan. A bypass area ratio is greater than about 6.0. The low pressure turbine section airfoil count to bypass area ratio is below about 170.
1 . A turbofan engine comprising:
a fan including a circumferential array of fan blades;
a compressor in fluid communication with the fan, the compressor including a second compressor section and a first compressor section, the second compressor section including a second compressor section inlet with a second compressor section inlet annulus area;
a fan duct including a fan duct annulus area outboard of the second compressor section inlet, wherein the ratio of the fan duct annulus area to the second compressor section inlet annulus area defines a bypass area ratio that is greater than 6.0;
a combustor in fluid communication with the compressor;
a shaft assembly having a first portion and a second portion;
a turbine in fluid communication with the combustor, the turbine having a two-stage first turbine section coupled to the first portion of the shaft assembly to drive the first compressor section, and a four-stage second turbine section coupled to the second portion of the shaft assembly to drive the fan, each of the second turbine section including blades and vanes, and a second turbine airfoil count defined as the numerical count of all of the blades and vanes in the second turbine section;
wherein a ratio of the second turbine airfoil count to the bypass area ratio is less than 150;
wherein the second turbine section further includes a maximum gas path radius and the fan blades include a maximum radius, and a ratio of the maximum gas path radius to the maximum radius of the fan blades is less than 0.55; and
wherein a hub-to-tip ratio (Ri:Ro) of the second turbine section is between 0.4 and 0.5 measured at the maximum Ro axial location in the second turbine section; and
a planetary gearbox coupled to the fan and rotatable by the second turbine section through the second portion of the shaft assembly to allow the second turbine to turn faster than the fan, the gearbox having a speed reduction ratio between 2:1 and 13:1 determined by the ratio of diameters within the gearbox.
2 . The turbofan engine as recited in claim 1 , wherein the bypass area ratio is greater than 8.0 and the sum of the plurality of stages of the second compressor section, first compressor section, second turbine section and first turbine section is no less than seventeen and no more than twenty one.
3 . The turbofan engine as recited in claim 2 , wherein the second compressor section is a four-stage second compressor.
4 . The turbofan engine as recited in claim 3 , wherein the first compressor section is a nine-stage first compressor.
5 . The turbofan engine as recited in claim 1 , wherein the bypass area ratio is greater than 8.0.
6 . The turbofan engine as recited in claim 5 , further comprising a fan case and vanes, the fan case encircling the fan and supported by the vanes, and wherein the second turbine section includes a plurality of blade stages interspersed with a plurality of vane stages, and each stage of the second turbine section includes a disk with a circumferential array of blades, each blade including an airfoil extending from an inner diameter to an outer diameter, wherein the inner diameter is associated with a platform and the outer diameter is associated with a shroud.
7 . The turbofan engine as recited in claim 1 , wherein the bypass area ratio is greater than 8.0 and further comprising:
an engine aft mount location configured to support an engine mount when the engine is mounted and to react at least a thrust load of the engine; and
an engine forward mount location configured to support an engine mount when the engine is mounted.
8 . The turbofan engine as recited in claim 7 , wherein the engine aft mount location engages with an engine thrust case.
9 . The turbofan engine as recited in claim 8 , wherein the engine aft mount location is located between the second turbine section and the first turbine section.
10 . The turbofan engine as recited in claim 9 , wherein the engine forward mount location engages with an intermediate case.
11 . A turbofan engine comprising:
a fan including a circumferential array of fan blades;
a compressor in fluid communication with the fan, the compressor including a second compressor section and a first compressor section, the second compressor section including a second compressor section inlet with a second compressor section inlet annulus area;
a fan duct including a fan duct annulus area outboard of the second compressor section inlet, wherein the ratio of the fan duct annulus area to the second compressor section inlet annulus area defines a bypass area ratio that is greater than 6.0;
a combustor in fluid communication with the compressor;
a turbine in fluid communication with the combustor, the turbine having a first turbine section and a second turbine section, wherein a hub-to-tip ratio (Ri:Ro) of the second turbine section is between 0.4 and 0.5 measured at the maximum Ro axial location in the second turbine section; and
an epicyclic gearbox coupled to the fan and rotatable by the second turbine section through the second portion of the shaft assembly to allow the second turbine section to turn faster than the fan, the gearbox having a speed reduction ratio between 2:1 and 13:1.
12 . The turbofan engine as recited in claim 11 , wherein the bypass area ratio is greater than 8.0 and wherein the sum of the plurality of stages of the second compressor section, first compressor section, second turbine section and first turbine section is no less than seventeen and no more than twenty one.
13 . The turbofan engine as recited in claim 12 , wherein the first turbine section is a two-stage first turbine.
14 . The turbofan engine as recited in claim 13 , wherein the second turbine section is a four-stage second turbine.
15 . The turbofan engine as recited in claim 11 , wherein the second compressor section is a four-stage second compressor.
16 . The turbofan engine as recited in claim 11 , wherein the first compressor section is a nine-stage first compressor.
17 . The turbofan engine as recited in claim 11 , wherein the bypass area ratio is greater than 8.0, the second turbine section is a four-stage second turbine and the first turbine section is a two-stage first turbine.
18 . The turbofan engine as recited in claim 17 , wherein the gearbox is a planetary gearbox, and further comprising a fan case and vanes, the fan case encircling the fan and supported by the vanes.
19 . The turbofan engine as recited in claim 11 , wherein the bypass area ratio is greater than 8.0 and the second turbine section includes a second turbine airfoil count defined as the numerical count of all of the blades and vanes in the second turbine section, wherein a ratio of the second turbine airfoil count to the bypass area ratio is less than 150.
20 . The turbofan engine as recited in claim 11 , wherein the bypass area ratio is greater than 8.0, the second turbine section further includes a maximum gas path radius, and the fan blades include a maximum radius, wherein a ratio of the maximum gas path radius to the maximum radius of the fan blades is less than 0.55.
21 . The turbofan engine as recited in claim 11 , wherein the bypass area ratio is greater than 8.0.
22 . The turbofan engine as recited in claim 11 , further comprising an engine intermediate case, including an engine forward mount location configured to support an engine mount when the engine is mounted, and an engine thrust case including an engine aft mount location configured to support an engine mount and react at least a thrust load when the engine is mounted.
23 . The turbofan engine as recited in claim 22 , wherein the engine aft mount location is located between the second turbine section and the first turbine section.
24 . A turbofan engine comprising:
a fan including a circumferential array of fan blades;
a compressor in fluid communication with the fan, the compressor including a first compressor section and a second compressor section;
a combustor in fluid communication with the compressor;
a turbine in fluid communication with the combustor, the turbine having a first turbine section and a second turbine section, the second turbine section driving the first compressor section and the fan, wherein:
the second turbine section includes no less than three stages and no more than five stages;
a hub-to-tip ratio (Ri:Ro) of the second turbine section is between 0.4 and 0.5 measured at the maximum Ro axial location in the second turbine section; and
a planetary gearbox coupled to the fan and rotatable by the second turbine section through the second portion of the shaft assembly to allow the second turbine to turn faster than the fan, the gearbox having a speed reduction ratio between 2:1 and 13:1.
25 . The turbofan engine as recited in claim 24 , wherein the second turbine section includes a second turbine airfoil count defined as the numerical count of all of the blades and vanes in the second turbine section, wherein a ratio of the second turbine airfoil count to the bypass area ratio is less than 150.
26 . The turbofan engine as recited in claim 24 , wherein the second turbine section further includes a maximum gas path radius and the fan blades include a maximum radius, and a ratio of the maximum gas path radius to the maximum radius of the fan blades is less than 0.55.
27 . The turbofan engine as recited in claim 24 , further comprising:
the second compressor section including a second compressor section inlet with a second compressor section inlet annulus area; and
a fan duct including a fan duct annulus area outboard of the second compressor section inlet, wherein the ratio of the fan duct annulus area to the second compressor section inlet annulus area defines a bypass area ratio that is greater than 6.0.
28 . The turbofan engine as recited in claim 24 , wherein the second turbine section is a four-stage second turbine and the first turbine is a two-stage first turbine.
29 . A turbine module for a geared turbofan engine comprising:
a plurality of stages configured to drive a compressor and cooperate with a gearbox to drive a fan upon connection of the turbine module with other sections of the geared turbofan engine; and
a core gaspath defined between an inboard boundary characterized by a radius Ri and an outboard boundary characterized by a radius Ro,
wherein a ratio of the radius Ri to the radius Ro (Ri:Ro) is between 0.4 and 0.5 measured at the maximum Ro axial location.
30 . The turbine module of claim 29 , wherein the plurality of stages is no less than three stages and no more than four stages.