Geared gas turbine engine with reduced fan noise

Patent metadata

Field	Value
Publication number	US-10107191-B2
Application number	US-201514964727-A
Country	US
Kind code	B2
Filing date	Dec 10, 2015
Priority date	Feb 29, 2012
Publication date	Oct 23, 2018
Grant date	Oct 23, 2018

Abstract

A fan section for a gas turbine engine according to an example of the present disclosure includes, among other things, a fan rotor having fan blades, and a plurality of fan exit guide vanes positioned downstream of the fan rotor. The fan rotor is configured to be driven through a gear reduction. A ratio of a number of fan exit guide vanes to a number of fan blades is defined. The fan exit guide vanes are provided with optimized sweep and optimized lean.

First claim

What is claimed is:

1. A gas turbine engine comprising:
a fan section having a plurality of fan blades, said fan section having a low fan pressure ratio of less than 1.45, wherein the low fan pressure ratio is measured across one of said fan blades alone;
a geared arrangement that drives said fan section;
a compressor section including a first compressor section and a second compressor section;
a turbine section including a first turbine section that drives said first compressor section, and a second turbine section that drives said second compressor section and said geared arrangement;
a plurality of exit guide vanes positioned downstream of said fan blades;
wherein a first ratio of a number of exit guide vanes to a number of fan blades being between 0.8 and 2.5; and
wherein said exit guide vanes are provided with sweep and lean,
wherein sweep means that: (a) an outer periphery of said exit guide vane is positioned to be downstream of a location of an inner periphery of said exit guide vane; and (b) a sweep angle is between 15 degrees and 30 degrees, and
wherein lean means that: (a) an outer periphery of said exit guide vane is positioned at a greater circumferential distance than an inner periphery of said exit guide vane in a direction of rotation of said fan blades; and (b) a lean angle is greater than or equal to 2 degrees and less than or equal to 15 degrees.

2. The gas turbine engine as set forth in claim 1 , wherein said sweep angle is greater than or equal to 20 degrees.

3. The gas turbine engine as set forth in claim 1 , wherein said fan section has 26 or fewer fan blades.

4. The gas turbine engine as set forth in claim 3 , wherein:
said geared arrangement has a gear reduction ratio greater than or equal to 2.3;
wherein the second 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 an exhaust nozzle; and
wherein the fan has a low corrected fan tip speed of less than 1150 ft/second.

5. The gas turbine engine as set forth in claim 4 , wherein said first ratio is greater than or equal to 2.0.

6. The gas turbine engine as set forth in claim 4 , wherein said first ratio is less than 2.0.

7. The gas turbine engine as set forth in claim 4 , wherein:
each of said exit guide vanes defines a hollow opening covered by an acoustic liner having a perforated face sheet;
said face sheet has a thickness, and a diameter of holes in said face sheet is selected to be less than or equal 0.3 of the thickness.

8. The gas turbine engine as set forth in claim 7 , wherein said holes in said face sheet are at least 5% of a surface area of said face sheet.

9. The gas turbine engine as set forth in claim 8 , wherein said diameter is less than or equal to 0.2 of the thickness.

10. The gas turbine engine as set forth in claim 1 , wherein:
each of said exit guide vanes defines a hollow opening covered by an acoustic liner having a perforated face sheet;
said face sheet has a thickness, and a diameter of holes in said face sheet is selected to be less than or equal 0.3 of the thickness.

11. The gas turbine engine as set forth in claim 10 , wherein said holes in said face sheet are at least 5% of a surface area of said face sheet.

12. The gas turbine engine as set forth in claim 11 , wherein said diameter is less than or equal to 0.2 of the thickness.

13. A gas turbine engine comprising:
a fan section having a plurality of fan blades, said fan section having a low fan pressure ratio of less than 1.45, wherein the low fan pressure ratio is measured across one of said fan blades alone;
a geared arrangement that drives said fan section;
a compressor section including a first compressor section and a second compressor section;
a turbine section including a first turbine section that drives said first compressor section, and a second turbine section that drives said second compressor section and said geared arrangement;
a plurality of exit guide vanes positioned downstream of said fan blades;
wherein a first ratio of a number of exit guide vanes to a number of fan blades being between 0.8 and 2.0; and
wherein said exit guide vanes are provided with sweep and lean,
wherein sweep means that: (a) an outer periphery of said exit guide vane is positioned to be downstream of a location of an inner periphery of said exit guide vane; and (b) a sweep angle is between 5 degrees and 35 degrees,
wherein lean means that: (a) an outer periphery of said exit guide vane is positioned at a greater circumferential distance than an inner periphery of said exit guide vane in a direction of rotation of said fan blades; and (b) a lean angle is greater than or equal to 2 degrees and less than or equal to 15 degrees.

14. The gas turbine engine as set forth in claim 13 , wherein said lean angle is greater than or equal to 2 degrees, and said sweep angle is greater than or equal to 20 degrees.

15. The gas turbine engine as set forth in claim 13 , wherein said fan section has 26 or fewer fan blades.

16. The gas turbine engine as set forth in claim 15 , wherein:
said geared arrangement has a gear reduction ratio greater than or equal to 2.3;
wherein the second 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 an exhaust nozzle; and
wherein the fan has a low corrected fan tip speed of less than 1150 ft/second.

17. The gas turbine engine as set forth in claim 16 , wherein said first ratio is greater than or equal to 2.0.

18. The gas turbine engine as set forth in claim 16 , wherein said first ratio is less than 2.0.

19. The gas turbine engine as set forth in claim 16 , wherein:
each of said exit guide vanes defines a hollow opening covered by an acoustic liner having a perforated face sheet;
said face sheet has a thickness, and a diameter of holes in said face sheet is selected to be less than or equal to 0.3 of the thickness; and
holes in said face sheet are at least 5% of a surface area of said face sheet.

20. The gas turbine engine as set forth in claim 19 , wherein said diameter is less than or equal to 0.2 of the thickness.

21. The gas turbine engine as set forth in claim 13 , wherein said geared arrangement is between said fan section and said second turbine section.

22. A gas turbine engine comprising:
a fan section having a plurality of fan blades;
a geared arrangement that drives said fan section;
a compressor section including a first compressor section and a second compressor section;
a turbine section including a first turbine section that drives said first compressor section, and a second turbine section that drives said second compressor section and said geared arrangement;
a plurality of exit guide vanes positioned downstream of said fan blades;
wherein a first ratio of a number of exit guide vanes to a number of fan blades being between 0.8 and 2.5;
wherein said exit guide vanes are provided with sweep and lean,
wherein sweep means that: (a) an outer periphery of said exit guide vane is positioned to be downstream of a location of an inner periphery of said exit guide vane; and (b) a sweep angle is between 5 degrees and 35 degrees,
wherein lean means that: (a) an outer periphery of said exit guide vane is positioned at a greater circumferential distance than an inner periphery of sa

Assignees

• United Technologies Corp

Inventors

• Gilson Jonthan
• Morin Burce L
• Reba Ramons A
• Topol David A
• Lord Wesley K

Classifications

• F05D2300/603

  Composites; e.g. fibre-reinforced · CPC title

• F05D2260/961

  by mistuning rotor blades or stator vanes with irregular interblade spacing, airfoil shape · CPC title

• F05D2260/96

  Preventing, counteracting or reducing vibration or noise · CPC title

• F05D2250/38

  angled, e.g. sweep angle · CPC title

• F05D2250/191

  perforated · CPC title