Turbine section of high bypass turbofan

Patent metadata

Field	Value
Publication number	US-11149650-B2
Application number	US-201816025038-A
Country	US
Kind code	B2
Filing date	Jul 2, 2018
Priority date	Aug 1, 2007
Publication date	Oct 19, 2021
Grant date	Oct 19, 2021

Abstract

A turbofan engine according to an example of the present disclosure includes, among other things, 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, 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, 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, an epicyclic transmission 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, wherein the second turbine section 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 0.6.

First claim

The invention claimed is:

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 first compressor section and a second compressor including four stages, the first compressor including a greater number of stages than the second compressor, 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 fan duct annulus area and the second compressor section inlet annulus area are established at a splitter that bounds the fan duct and the second compressor section inlet, and 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 or equal to 8.0;
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; and
an epicyclic transmission 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;
wherein the second turbine airfoil count is below 1600;
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 greater than 0.55 and less than or equal to 0.65.

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 epicyclic transmission is a planetary gearbox having a speed reduction ratio between 2:1 and 13:1 determined by the ratio of diameters within the planetary gearbox.

5. The turbofan engine as recited in claim 4 , wherein the ratio of the maximum gas path radius to the maximum radius of the fan blades is less than 0.6.

6. The turbofan engine as recited in claim 3 , further comprising:
an engine aft mount location configured 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.

7. The turbofan engine as recited in claim 6 , wherein the engine forward mount location is axially proximate to the epicyclic transmission.

8. The turbofan engine as recited in claim 7 , wherein the engine forward mount location engages with an intermediate case.

9. The turbofan engine as recited in claim 6 , wherein the engine aft mount location engages with an engine thrust case.

10. The turbofan engine as recited in claim 9 , wherein the engine aft mount location is located between the second turbine section and the first turbine section.

11. The turbofan engine as recited in claim 10 wherein the engine aft mount location is located between the second turbine section and the first turbine section.

12. 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.

13. The turbofan engine as recited in claim 12 , wherein in at least one stage the shroud is integral with the airfoil.

14. The turbofan engine as recited in claim 13 , wherein the shroud includes outboard sealing ridges configured to seal with abradable seals.

15. The turbofan engine as recited in claim 14 , wherein the abradable seals include honeycomb.

16. The turbofan engine as recited in claim 15 , further comprising a case associated with the second turbine section, wherein the abradable seals are fixed to the case.

17. The turbofan engine as recited in claim 16 , 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.

18. The turbofan engine as recited in claim 17 , further comprising a stationary blade outer air seal, and a rotational gap between a tip of the airfoil and the stationary blade outer air seal.

19. The turbofan engine as recited in claim 18 , 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.

20. The turbofan engine as recited in claim 19 , further comprising a case associated with the second turbine section, wherein the shroud is fixed to the case.

21. 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.

22. The turbofan engine as recited in claim 21 , wherein the bypass area ratio is less than or equal to 20, the epicyclic transmission is a planetary gearbox having a speed reduction ratio between 2:1 and 13:1 determined by the ratio of diameters within the planetary gearbox, and the hub-to-tip ratio (Ri:Ro) is between 0.42-0.48.

23. The turbofan engine as recited in claim 22 , wherein the ratio of the maximum gas path radius to the maximum radius of the fan blades is greater than 0.55 and less than 0.6 and the ratio of the second turbine airfoil count to the bypass area ratio is between 120 and 140.

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 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 fan duct annulus area and the second compressor section inlet annulus area are established at a splitter that bounds the fan duct and the second compressor section inlet, and 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 firs

Assignees

• Raytheon Tech Corp

Inventors

• Adams Paul R
• Magge Shankar S
• Staubach Joseph B
• Lord Wesley K
• Schwarz Frederick M
• Suciu Gabriel L

Classifications

• F05B2250/283

  Honeycomb · CPC title

• F04D29/324

  Blades · CPC title

• F04D29/545

  Ducts · CPC title

• F02C7/36 Primary

  Power transmission arrangements between the different shafts of the gas turbine plant, or between the gas-turbine plant and the power user ({F02C3/107 - F02C3/13 and} F02C7/32 take precedence; couplings for transmitting rotation F16D; gearing in general F16H) · CPC title

• F05D2220/32

  in gas turbines · CPC title