Turbine section of high bypass turbofan
A turbofan engine according to an example of the present disclosure includes, among other things, 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 fan rotor and a circumferential array of fan blades;
a compressor in fluid communication with the fan, the compressor including a four-stage second compressor section and a nine-stage 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 between 8.0 and 20.0;
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 second turbine section coupled to the second portion of the shaft assembly to drive the fan, the second turbine section being a four-stage turbine section, 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; 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 section 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, and wherein the planetary gearbox is intermediate the second compressor section and the second portion of the shaft assembly such that the fan rotor and the second compressor section are rotatable at a common speed;
wherein a ratio of the second turbine airfoil count to the bypass area ratio is less than 150, and
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.
2. The turbofan engine as recited in claim 1 , further comprising a fan case and vanes, the fan case encircling the fan and supported by the vanes.
3. The turbofan engine as recited in claim 2 , wherein the fan is a single fan, and each fan blade includes a platform and an outboard end having a free tip.
4. The turbofan engine as recited in claim 3 , wherein the gearbox carries a plurality of gears associated with journals.
5. The turbofan engine as recited in claim 3 , further comprising:
an engine aft mount location to support an engine mount when the engine is mounted and react at least a thrust load of the engine; and
an engine forward mount location.
6. The turbofan engine as recited in claim 5 , wherein the engine forward mount location is axially proximate to the gearbox.
7. The turbofan engine as recited in claim 6 , wherein the engine forward mount location engages with an intermediate case.
8. The turbofan engine as recited in claim 5 , 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 1 , 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.
11. The turbofan engine as recited in claim 10 , wherein in at least one stage the shroud is integral with the airfoil.
12. The turbofan engine as recited in claim 11 , wherein the shroud includes outboard sealing ridges configured to seal with abradable seals.
13. The turbofan engine as recited in claim 12 , wherein the abradable seals include honeycomb.
14. The turbofan engine as recited in claim 13 , further comprising a case associated with the second turbine section, wherein the abradable seals are fixed to the case.
15. The turbofan engine as recited in claim 14 , wherein 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 unshrouded.
16. The turbofan engine as recited in claim 14 , further comprising a stationary blade outer air seal, and a rotational gap between the tip and the stationary blade outer air seal.
17. The turbofan engine as recited in claim 16 , wherein each of the plurality of vane stages includes a vane, each vane 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.
18. The turbofan engine as recited in claim 17 , further comprising a case associated with the second turbine section, wherein the shroud is fixed to the case.
19. The turbofan engine as recited in claim 18 , wherein each platform carries a seal.
20. The turbofan engine as recited in claim 1 , 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.
21. A turbofan engine comprising:
a fan including a fan rotor and 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 between 8.0 and 20.0;
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 first turbine section coupled to the first portion of the shaft assembly to drive the first compressor section, and a second turbine section coupled to the second portion of the shaft assembly to drive the fan, 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; 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 section 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, and wherein the planetary gearbox is intermediate the second compressor section and the second portion of the shaft assembly such that the fan rotor and the second compressor section are rotatable at a common speed;
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.
22. The turbofan engine as recited in claim 21 , wherein the first turbine is a two-stage first turbine and the second turbine section is a four-stage second turbine.
23. The turbofan engine as recited in claim 22 , wherein the first compressor is a nine-stage first compressor.
24. The turbofan engine as recited in claim 23 , wherein the second compressor is a four-stage second compressor.
25. The turbofan engine as recited in claim 23 , further comprising a fan case and vanes, the fan case encircling the fan and supported by the vanes, and wherein the gearbox carries a plurality of gears associated with journals.
26. The turbofan engine as recited in claim 23 , further comprising an engine intermediate case, including an engine forward mount location proximate to the gearbox and 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, wherein the engine aft mount location is located between the second turbine section and the first turbine section.
27. The turbofan engine as recited in claim 21 , wherein:
the fan is a single fan;
the ratio of the maximum gas path radius to the maximum radius of the fan blades is between 0.35 and 0.50; and
the hub-to-tip ratio (Ri:Ro) is between 0.42 and 0.48.