Turbine section of high bypass turbofan

Patent metadata

Field	Value
Publication number	US-11242805-B2
Application number	US-201816025022-A
Country	US
Kind code	B2
Filing date	Jul 2, 2018
Priority date	Aug 1, 2007
Publication date	Feb 8, 2022
Grant date	Feb 8, 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 first compressor section and a second compressor section, the second compressor section including a second compressor section inlet with a compressor inlet annulus area, a fan duct including a fan duct annulus area outboard of the second compressor section inlet, and a turbine having a first turbine section driving the first compressor section, a second turbine section driving the fan through an epicyclic gearbox, the second turbine section including blades and vanes, and wherein the second 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 including an array of fan blades rotatable about an engine axis;
a compressor including a first compressor section and a second compressor section, the second compressor section including a second compressor section inlet with a compressor 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 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 first turbine section driving the first compressor section, and a second turbine section driving the fan through an epicyclic gearbox, a mid-turbine frame between the first turbine section and the second turbine section, the second turbine section being a two-stage to four-stage turbine, 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 between 10 and 150;
wherein the second 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 second turbine section is greater than 0.5, measured at the maximum Ro axial location in the second turbine section.

2. The turbofan engine as recited in claim 1 , wherein the fan is a single fan, and the array of fan blades have a fixed stagger angle.

3. The turbofan engine as recited in claim 2 , 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.

4. The turbofan engine as recited in claim 3 , wherein the engine forward mount location engages with an intermediate case.

5. The turbofan engine as recited in claim 3 , wherein the engine aft mount location engages with the mid-turbine frame.

6. The turbofan engine as recited in claim 5 , wherein the mid-turbine frame supports at least one bearing, and includes a plurality of airfoils distributed in a core flow path.

7. The turbofan engine as recited in claim 1 , wherein the first turbine section is a two-stage turbine, the second turbine section is a three-stage or four-stage turbine, the second turbine 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.

8. The turbofan engine as recited in claim 7 , wherein the epicyclic gearbox is a planetary gear system.

9. 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 less than 0.50.

10. The turbofan engine as recited in claim 9 , 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 second turbine section.

11. The turbofan engine as recited in claim 10 , wherein the second 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.

12. The turbofan engine as recited in claim 10 , wherein the second 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.

13. The turbofan engine as recited in claim 10 , wherein the array of fan blades have a fixed stagger angle.

14. The turbofan engine as recited in claim 10 , wherein the gearbox is located aft of the first compressor section.

15. The turbofan engine as recited in claim 10 , 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.

16. The turbofan engine as recited in claim 15 , wherein the fan nacelle defines an engine inlet, the variable area fan nozzle defines a bypass outlet, and a pressure ratio defined by the engine inlet and the bypass outlet being less than or equal to 1.4.

17. The turbofan engine as recited in claim 16 , wherein the second turbine section is a three-stage or a four-stage turbine, the hub-to-tip ratio (Ri:Ro) is between 0.55 and 0.65, measured at the maximum Ro axial location in the second turbine section.

18. The turbofan engine as recited in claim 17 , wherein the array of fan blades comprise a composite material.

19. 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, the second compressor section including a second compressor section inlet with a compressor 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 compressor inlet annulus area defines a bypass area ratio between 8.0 and 20.0;
a turbine having a first turbine section driving the first compressor section, and a second turbine section driving the fan through an epicyclic gearbox, a mid-turbine frame between the first turbine section and the second turbine section, the second turbine section being a two-stage to four-stage turbine, 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 between 10 and 150;
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 equal to or greater than 0.35, and is less than 0.6; and
wherein a hub-to-tip ratio (Ri:Ro) of the second turbine section is greater than 0.5, and is less than or equal to 0.7, measured at the maximum Ro axial location in the second turbine section.

20. The turbofan engine as recited in claim 19 , wherein the first turbine section is a two-stage turbine and the second turbine section is a four-stage turbine.

21. The turbofan engine as recited in claim 20 , wherein the first compressor section is a nine-stage compressor, and the second compressor section is a four-stage compressor.

22. The turbofan engine as recited in claim 21 , further comprising an engine intermediate case, an engine forward mount location proximate to the gearbox and supporting an engine mount when the engine is mounted, and an engine thrust case including an engine aft mount location supporting an engine mount and to react at least a t

Assignees

• Raytheon Tech Corp

Inventors

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

Classifications

• F01D7/00

  Rotors with blades adjustable in operation; Control thereof (for reversing F01D1/30) · CPC title

• F05D2250/80

  Size or power range of the machines · CPC title

• F04D29/324

  Blades · CPC title

• F02C3/107

  with two or more rotors connected by power transmission · CPC title

• F02C9/18

  by bleeding, bypassing or acting on variable working fluid interconnections between turbines or compressors or their stages {(F02C3/113 takes precedence)} · CPC title