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, and a core including a core flowpath, the fan configured to communicate a portion of air into the core flowpath and a portion of air into a bypass flowpath;
a compressor in fluid communication with the fan, the compressor including a low pressure compressor section and a high pressure compressor section, the low pressure compressor section including a low pressure compressor section inlet with a low pressure compressor section inlet annulus area;
a fan duct including a fan duct annulus area outboard of the low pressure compressor section inlet, wherein the ratio of the fan duct annulus area to the low pressure compressor section inlet annulus area defines a bypass area ratio that is greater than 8.0;
a combustor in fluid communication with the compressor;
a shaft assembly having a high portion and a low portion;
a turbine in fluid communication with the combustor, the turbine having a high pressure turbine section coupled to the high portion of the shaft assembly to drive the high pressure compressor section, and a low pressure turbine section coupled to the low portion of the shaft assembly to drive the fan, the low pressure turbine section including blades and vanes, and a low pressure turbine airfoil count defined as the numerical count of all of the blades and vanes in the low pressure turbine section;
wherein a ratio of the low pressure turbine airfoil count to the bypass area ratio is less than 150;
wherein the low pressure 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
an epicyclic gearbox coupled to the fan and rotatable by the low pressure turbine section through the low portion of the shaft assembly to allow the low pressure 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;
wherein each of the low pressure compressor section, high pressure compressor section, low pressure turbine section and high pressure turbine section include a plurality of stages, wherein:
each of the plurality of stages of the low pressure turbine includes a disk, and a circumferential array of blades extend from the disk, each of the blades including an airfoil extending from a platform at an inner diameter, to an outer diameter shroud, the shroud having an outboard seal which seals with an abradable fixed seal.
2 . The turbofan engine as recited in claim 1 , wherein the sum of the plurality of stages of the low pressure compressor section, high pressure compressor section, low pressure turbine section and high pressure turbine section is no less than seventeen and no more than twenty one.
3 . The turbofan engine as recited in claim 2 , wherein the high pressure turbine is a two-stage high pressure turbine.
4 . The turbofan engine as recited in claim 3 , wherein the plurality of stages of the low pressure turbine is no less than two and no more than five.
5 . The turbofan engine as recited in claim 4 , wherein the low pressure turbine is a four-stage low pressure turbine.
6 . The turbofan engine as recited in claim 5 , wherein the sum of the plurality of stages of the low pressure compressor section, high pressure compressor section, low pressure turbine section and high pressure turbine section is twenty one.
7 . The turbofan engine as recited in claim 4 , wherein the sum of the plurality of stages of the low pressure compressor section, high pressure compressor section, low pressure turbine section and high pressure turbine section is no less than eighteen and no more than twenty.
8 . The turbofan engine as recited in claim 7 , wherein the low pressure compressor is a four-stage low pressure compressor.
9 . The turbofan engine as recited in claim 7 , wherein the high pressure compressor is a nine-stage high pressure compressor.
10 . The turbofan engine as recited in claim 1 , wherein the low pressure turbine is a four-stage low pressure turbine and the high pressure turbine is a two-stage high pressure turbine.
11 . The turbofan engine as recited in claim 10 , further comprising a fan case and vanes, the fan case encircling the fan and supported by the vanes.
12 . The turbofan engine as recited in claim 11 , wherein the gearbox is a planetary gearbox.
13 . The turbofan engine as recited in claim 12 , wherein the gearbox carries a plurality of gears associated with journals.
14 . The turbofan engine as recited in claim 12 , 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.
15 . The turbofan engine as recited in claim 11 , wherein the gearbox is a star gearbox and includes a sun gear, a ring gear, and a carrier that carries a plurality of star gears positioned between and enmeshed with the sun gear and ring gear, the star gears having associated journals.
16 . The turbofan engine as recited in claim 11 , wherein a plurality of the fan blades are mounted in a fixed relationship to a hub.
17 . The turbofan engine as recited in claim 1 , wherein the low pressure turbine includes a plurality of rows of the blades, and a row of the vanes intermediate a plurality of the rows of the blades.
18 . A turbofan engine comprising:
a fan including a circumferential array of fan blades, and a core including a core flowpath, the fan configured to communicate a portion of air into the core flowpath and a portion of air into a bypass flowpath;
a compressor in fluid communication with the fan, the compressor including a low pressure compressor section and a high pressure compressor section, the low pressure compressor section including a low pressure compressor section inlet with a low pressure compressor section inlet annulus area;
a fan duct including a fan duct annulus area outboard of the low pressure compressor section inlet, wherein the ratio of the fan duct annulus area to the low pressure compressor section inlet annulus area defines a bypass area ratio that is greater than 8.0;
a combustor in fluid communication with the compressor;
a shaft assembly having a high portion and a low portion;
a turbine in fluid communication with the combustor, the turbine having a high pressure turbine section coupled to the high portion of the shaft assembly to drive the high pressure compressor section, and a low pressure turbine section coupled to the low portion of the shaft assembly to drive the fan, the low pressure turbine section including blades and vanes, and a low pressure turbine airfoil count defined as the numerical count of all of the blades and vanes in the low pressure turbine section;
wherein a ratio of the low pressure turbine airfoil count to the bypass area ratio is less than 150;
wherein the low pressure 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;
wherein each of the low pressure compressor section, high pressure compressor section, low pressure turbine section and high pressure turbine section include a plurality of stages, and the sum of such stages being no less than seventeen and no more than twenty one; and
an epicyclic planetary gearbox coupled to the fan and rotatable by the low pressure turbine section through the low portion of the shaft assembly to allow the low pressure 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.
19 . The turbofan engine as recited in claim 18 , wherein the sum of the plurality of stages of the low pressure compressor section, high pressure compressor section, low pressure turbine section and high pressure turbine section is no less than eighteen and no more than twenty.
20 . The turbofan engine as recited in claim 19 , wherein the sum of the plurality of stages of the low pressure compressor section, high pressure compressor section, low pressure turbine section and high pressure turbine section is no less than eighteen and no more than nineteen.
21 . The turbofan engine as recited in claim 20 , wherein the high pressure turbine is a two-stage high pressure turbine.
22 . The turbofan engine as recited in claim 21 , wherein the low pressure turbine is a four-stage low pressure turbine.
23 . The turbofan engine as recited in claim 22 , 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.
24 . The turbofan engine as recited in claim 18 , wherein the low pressure turbine is a four-stage low pressure turbine and the high pressure turbine is a two-stage high pressure turbine.
25 . The turbofan engine as recited in claim 24 , wherein the low pressure compressor is a four-stage low pressure compressor.
26 . The turbofan engine as recited in claim 25 , wherein the high pressure compressor is a nine-stage high pressure compressor.
27 . The turbofan engine as recited in claim 24 , wherein the blades in the low pressure turbine include an airfoil extending from an inner diameter platform to an outer diameter shroud, and the shroud has an outboard seal which seals with an abradable fixed seal.
28 . The turbofan engine as recited in claim 18 , further comprising a fan case and vanes, the fan case encircling the fan and supported by the vanes, and wherein the gearbox is a planetary gearbox.
29 . The turbofan engine as recited in claim 28 , wherein each fan blade includes a platform and an outboard end having a free tip, and is fixed in position between the platform and the free tip.
30 . The turbofan engine as recited in claim 29 , 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.