Turbofan engine front section

Patent metadata

Field	Value
Publication number	US-11053816-B2
Application number	US-201615007263-A
Country	US
Kind code	B2
Filing date	Jan 27, 2016
Priority date	May 9, 2013
Publication date	Jul 6, 2021
Grant date	Jul 6, 2021

Abstract

A front section for a gas turbine engine according to an example of the present disclosure includes, among other things, a fan section including a fan hub. The fan hub includes a hub diameter supporting a plurality of fan blades including a tip diameter. A transitional entrance passage is configured to communicate flow between the fan section and a compressor section. The transitional entrance passage includes an inlet disposed at an inlet diameter and an outlet to the compressor section disposed at an outlet diameter. A method of designing a gas turbine engine is also disclosed.

First claim

What is claimed is:

1. A front section for a gas turbine engine comprising:
a fan section including a fan hub, the fan hub including a hub diameter supporting a plurality of fan blades including a tip diameter, a ratio of the hub diameter to the tip diameter being between 0.24 and 0.36; and
a transitional entrance passage configured to communicate flow between the fan section and a compressor section, the transitional entrance passage including an inlet disposed at an inlet diameter and an outlet disposed at an outlet diameter, the outlet adjacent to the compressor section, and a ratio of the hub diameter to the inlet diameter being between 0.70 and 0.90; and
wherein the outlet is axially forward of a forwardmost stage of the compressor section relative to an engine longitudinal axis, and a ratio of the inlet diameter to the outlet diameter is between 1.10 and 1.64.

2. The front section as recited in claim 1 , wherein the plurality of fan blades is less than twenty (20) fan blades.

3. The front section as recited in claim 1 , wherein a pressure ratio across the fan section is less than 1.5 across the the plurality of fan blades.

4. The front section as recited in claim 1 , wherein the compressor section includes a multi-stage low pressure compressor.

5. A gas turbine engine comprising:
a fan section including a fan hub, the fan hub including a hub diameter supporting a plurality of fan blades including a tip diameter, a ratio of the hub diameter to the tip diameter being between 0.24 and 0.36, and an outer housing surrounding the fan blades to establish a bypass duct;
a compressor section including a low pressure compressor and a high pressure compressor;
a turbine section including a fan drive turbine configured to drive the fan section and the low pressure compressor, the fan drive turbine having a greater number of stages than the low pressure compressor, and the fan drive turbine having fewer stages than the high pressure compressor; and
a transitional entrance passage coupling the fan section and the compressor section, the transitional entrance passage including an inlet disposed at an inlet diameter and an outlet disposed at an outlet diameter, the inlet adjacent to the fan section, the outlet adjacent to the compressor section, and a ratio of the hub diameter to the inlet diameter being between 0.70 and 0.90; and
wherein the outlet is axially forward of a forwardmost stage of the compressor section relative to an engine longitudinal axis, and a ratio of the inlet diameter to the outlet diameter is between 1.10 and 1.64.

6. The gas turbine engine as recited in claim 5 , wherein the low pressure compressor is a multi-stage compressor.

7. The gas turbine engine as recited in claim 6 , comprising a geared architecture configured to rotate the fan hub at a lower relative speed than the fan drive turbine.

8. The gas turbine engine as recited in claim 5 , wherein the plurality of fan blades is less than twenty (20) fan blades.

9. The gas turbine engine as recited in claim 8 , wherein the low pressure compressor is a three stage compressor.

10. The gas turbine engine as recited in claim 9 , wherein the turbine section includes a high pressure turbine configured to drive the high pressure compressor, the high pressure turbine including at least two stages.

11. The gas turbine engine as recited in claim 6 , wherein the fan section includes an outer housing that surrounds the plurality of fan blades, the turbine section includes a high pressure turbine configured to drive the high pressure compressor, the high pressure turbine including at least two stages, the fan drive turbine having a greater number of stages than the high pressure turbine, the low pressure compressor includes three compressor stages, the high pressure compressor having a greater number of stages than the low pressure compressor, the plurality of fan blades is less than twenty (20) fan blades, and a ratio between the number of fan blades and the number of turbine rotors of the fan drive turbine is between 3.3 and 8.6.

12. The gas turbine engine as recited in claim 11 , wherein the fan drive turbine includes three turbine rotors.

13. The gas turbine engine as recited in claim 12 , wherein a pressure ratio across the fan section is less than 1.45 across the the plurality of fan blades.

14. The gas turbine engine as recited in claim 13 , comprising a geared architecture including an epicyclic gear train configured to rotate the fan hub at a lower relative speed than the fan drive turbine.

15. The gas turbine engine as recited in claim 14 , wherein the geared architecture is a planetary gear system.

16. The gas turbine engine as recited in claim 11 , wherein the fan drive turbine includes five turbine rotors, a pressure ratio across the fan section is less than 1.45 across the plurality of fan blades, and further comprising:
a geared architecture including an epicyclic gear train configured to rotate the fan hub at a lower relative speed than the fan drive turbine, wherein the geared architecture is a planetary gear system.

Assignees

• United Technologies Corp
• Raytheon Tech Corp

Inventors

• Schwarz Frederick M
• Sheridan William G

Classifications

• F02K3/06 Primary

  with front fan · CPC title

• F01D5/12

  Blades ({specially adapted for radial flow machines or engines F01D5/04}; blade roots F01D5/30; rotors with blades adjustable in operation F01D7/00; stator blades F01D9/02) · CPC title

• F05D2230/60

  Assembly methods · CPC title

• F05D2220/32

  in gas turbines · CPC title

• F05D2240/24

  for turbines · CPC title