Turbine section of high bypass turbofan

Patent metadata

Field	Value
Publication number	US-11486311-B2
Application number	US-202117230326-A
Country	US
Kind code	B2
Filing date	Apr 14, 2021
Priority date	Aug 1, 2007
Publication date	Nov 1, 2022
Grant date	Nov 1, 2022

Abstract

A turbofan engine according to an example of the present disclosure includes, among other things, a fan including an array of fan blades rotatable about an engine axis, a compressor including a high pressure compressor section and a low pressure compressor section, the low pressure compressor section including a low pressure compressor section inlet with a low pressure compressor inlet annulus area, a fan duct including a fan duct annulus area outboard of the low pressure compressor section inlet, and a turbine having a high pressure turbine section and a low pressure turbine section driving the fan through a speed reduction mechanism, wherein the low pressure turbine section defines a maximum gas path radius and the fan blades define 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

What is claimed is:

1. A turbofan engine comprising:
a fan section including a single-stage fan and a fan case encircling the fan to establish a fan duct, the fan including an array of fan blades rotatable about an engine axis;
a compressor including a high pressure compressor section and a low pressure compressor section having three stages, the low pressure compressor section including a low pressure compressor section inlet with a compressor inlet annulus area;
wherein the fan duct includes a fan duct annulus area outboard of the low pressure compressor section inlet, the fan duct annulus area and the low pressure compressor section inlet annulus area are established at a splitter that bounds the fan duct and the low pressure compressor section inlet, and a ratio of the fan duct annulus area to the compressor inlet annulus area defines a bypass area ratio between 8.0 and 20.0;
a turbine having a two-stage high pressure turbine section driving the high pressure compressor section, and a low pressure turbine section driving the fan through a speed reduction mechanism, the low pressure turbine section being a three-stage to four-stage turbine, 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 between 10 and 150;
wherein the low pressure turbine section defines a maximum gas path radius and the fan blades define 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; and
wherein a hub-to-tip ratio (Ri:Ro) of the low pressure turbine section is greater than 0.5, measured at the maximum Ro axial location in the low pressure turbine section.

2. The turbofan engine as recited in claim 1 , wherein the low pressure turbine section includes an inlet, an outlet, and a pressure ratio of greater than 5, the pressure ratio being pressure measured prior to the inlet as related to pressure at the outlet prior to any exhaust nozzle, and further comprising:
a mid-turbine frame between the high pressure turbine section and the low pressure turbine section, wherein the mid-turbine frame supports at least one bearing, and the mid-turbine frame includes a plurality of airfoils distributed in a core flow path.

3. The turbofan engine as recited in claim 2 , wherein the speed reduction mechanism is an epicyclic transmission.

4. The turbofan engine as recited in claim 3 , wherein the ratio of the maximum gas path radius to the maximum radius of the fan blades is equal to or greater than 0.35.

5. The turbofan engine as recited in claim 3 , further comprising a fan pressure ratio of less than 1.45 across the fan blade alone at cruise at 0.8 Mach and 35,000 ft.

6. The turbofan engine as recited in claim 5 , wherein the hub-to-tip ratio (Ri:Ro) is less than or equal to 0.7, measured at the maximum Ro axial location in the low pressure turbine section.

7. The turbofan engine as recited in claim 6 , wherein the array of fan blades have a fixed stagger angle.

8. The turbofan engine as recited in claim 7 , wherein the ratio of the maximum gas path radius to the maximum radius of the fan blades is equal to or greater than 0.35, and is less than 0.50, and wherein the low pressure turbine airfoil count is less than 1700.

9. The turbofan engine as recited in claim 8 , wherein the low pressure turbine airfoil count is less than 1000.

10. The turbofan engine as recited in claim 8 , wherein the low pressure turbine section includes a plurality of blade stages interspersed with a plurality of vane stages, with an aftmost one of the blade stages including shrouded blades.

11. The turbofan engine as recited in claim 8 , wherein the low pressure turbine section includes a plurality of blade stages interspersed with a plurality of vane stages, with an aftmost one of the blade stages including unshrouded blades.

12. The turbofan engine as recited in claim 2 , further comprising:
a fan nacelle and a core nacelle, the fan nacelle at least partially surrounding the core nacelle; and
a variable area fan nozzle in communication with the fan duct, and defining a fan nozzle exit area between the fan nacelle and the core nacelle, the variable area fan nozzle moveable to change the fan nozzle exit area.

13. The turbofan engine as recited in claim 12 , further comprising a fan pressure ratio of less than or equal to 1.4 across the fan blade alone at cruise at 0.8 Mach and 35,000 ft.

14. The turbofan engine as recited in claim 13 , wherein the hub-to-tip ratio (Ri:Ro) is less than or equal to 0.7, measured at the maximum Ro axial location in the low pressure turbine section.

15. The turbofan engine as recited in claim 13 , wherein the fan nacelle includes a first fan nacelle section and a second fan nacelle section, and the variable area fan nozzle establishes an auxiliary port axially aft of the splitter relative to the engine axis that communicates bypass flow in response to movement between the first fan nacelle section and the second fan nacelle section.

16. The turbofan engine as recited in claim 12 , wherein the array of fan blades comprise a composite material.

17. The turbofan engine as recited in claim 12 , wherein the array of fan blades have a fixed stagger angle.

18. The turbofan engine as recited in claim 12 , wherein the fan section includes a pitch change mechanism that causes each of the fan blades to rotate about a respective fan blade axis between a first position and a second position to vary a pitch of the respective fan blade.

19. A turbofan engine comprising:
a fan section including a single-stage fan and a fan case encircling the fan to establish a fan duct, the fan including a circumferential array of fan blades rotatable about an engine axis;
a compressor in fluid communication with the fan, the compressor including a high pressure compressor section and a low pressure compressor section having three stages, the low pressure compressor section including a low pressure compressor section inlet with a compressor inlet annulus area;
wherein the fan duct includes a fan duct annulus area outboard of the low pressure compressor section inlet, the fan duct annulus area and the low pressure compressor section inlet annulus area are established at a splitter that bounds the fan duct and the low pressure compressor section inlet, and the ratio of the fan duct annulus area to the compressor inlet annulus area defines a bypass area ratio between 8.0 and 20.0;
a turbine having a two-stage high pressure turbine section driving the high pressure compressor section, and a low pressure turbine section driving the fan through a speed reduction mechanism, the low pressure turbine section being a three-stage to four-stage turbine, 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 between 10 and 150;
wherein the low pressure 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 equal to or greater than 0.35, and is less than 0.6; and
wherein a hub-to-tip ratio (Ri:Ro) of the low pressure turbine section is greater than 0.5, and

Assignees

• Raytheon Tech Corp

Inventors

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

Classifications

• F02K3/06

  with front fan · CPC title

• F01D5/06

  Rotors for more than one axial stage, e.g. of drum or multiple disc type; Details thereof, e.g. shafts, shaft connections {(F01D5/022, F01D5/023 take precedence)} · CPC title

• F05D2250/80

  Size or power range of the machines · CPC title

• F02C7/20

  Mounting or supporting of plant; Accommodating heat expansion or creep · CPC title

• F05D2220/32

  in gas turbines · CPC title