Compressor flowpath

Patent metadata

Field	Value
Publication number	US-11428242-B2
Application number	US-201916715528-A
Country	US
Kind code	B2
Filing date	Dec 16, 2019
Priority date	Jan 31, 2012
Publication date	Aug 30, 2022
Grant date	Aug 30, 2022

Abstract

A compressor section for a gas turbine engine according to an example of the present disclosure includes, among other things, a low pressure compressor including a plurality of rotor blades arranged about an axis, a high pressure compressor, and a core flowpath passing through the low pressure compressor. The core flowpath at the low pressure compressor defines an inner diameter and an outer diameter relative to the axis. The outer diameter has a slope angle relative to the axis.

First claim

What is claimed is:

1. A gas turbine engine comprising:
a fan section including a fan surrounded by an outer housing;
a compressor section including a low pressure compressor having three stages defining an axis, a high pressure compressor including a greater number of stages than the low pressure compressor, and a core flowpath passing through the low pressure compressor;
a turbine section including a high pressure turbine having two stages and a low pressure turbine having five stages, the high pressure turbine driving the high pressure compressor, and the low pressure turbine driving the low pressure compressor and the fan section; and
wherein the core flowpath in the low pressure compressor has an inner diameter and an outer diameter relative to the axis, the outer diameter has a slope angle that is between 10 degrees and 15 degrees relative to the axis, and wherein the inner diameter of the core flowpath increases through the low pressure compressor along a fluid flow direction; and
wherein the fan delivers air into a bypass duct, and a portion of air into the compressor section, with a bypass ratio defined as the volume of air delivered into the bypass duct compared to the volume of air delivered into the compressor section, and the bypass ratio being greater than 10.

2. The gas turbine engine of claim 1 , wherein the slope angle slopes toward the axis along a fluid flow direction of the core flowpath.

3. The gas turbine engine of claim 2 , wherein the outer diameter is defined by one of a plurality of rotor blades of the low pressure compressor.

4. The gas turbine of claim 1 , wherein the low pressure compressor includes at least one variable vane situated in the core flowpath.

5. The gas turbine engine of claim 1 , wherein the low pressure compressor includes an exit guide vane located in a low pressure compressor outlet section of the core flowpath.

6. The gas turbine engine of claim 1 , wherein the gas turbine engine is a two-spool turbofan.

7. The gas turbine engine of claim 1 , wherein a pressure ratio across the fan section is less than 1.6 across the fan blade alone.

8. The gas turbine engine of claim 7 , wherein the low pressure turbine includes an inlet, an outlet, a pressure ratio greater than 5, the pressure ratio of the low pressure turbine being pressure measured prior to the inlet as related to pressure at the outlet prior to an exhaust nozzle.

9. The gas turbine engine of claim 8 , wherein:
the gas turbine engine is a two-spool turbofan;
the slope angle slopes toward the axis along a fluid flow direction of the core flowpath;
the low pressure turbine includes a greater number of stages than the low pressure compressor; and
the low pressure compressor includes a greater number of stages than the high pressure turbine.

10. The gas turbine engine of claim 9 , wherein the fan section has only a single fan stage comprising the fan.

11. The gas turbine engine of claim 9 , wherein the outer diameter is defined by one of a plurality of rotor blades of the low pressure compressor.

12. The gas turbine engine of claim 9 , wherein the low pressure compressor includes an exit guide vane located in a low pressure compressor outlet section of the core flowpath.

13. The gas turbine of claim 9 , wherein the low pressure compressor includes at least one variable vane situated in the core flowpath.

14. A gas turbine engine comprising:
a fan section including a fan surrounded by an outer housing;
a compressor section including a three-stage low pressure compressor defining an axis, a high pressure compressor including a greater number of stages than the low pressure compressor, and a core flowpath passing through the low pressure compressor;
a turbine section including a two-stage high pressure turbine and a low pressure turbine including five stages, the high pressure turbine driving the high pressure compressor, the low pressure turbine driving the low pressure compressor and driving the fan section through a geared architecture, the low pressure turbine including a greater number of stages than the low pressure compressor, the low pressure compressor including a greater number of stages than the high pressure turbine; and
wherein the core flowpath within the low pressure compressor has an inner diameter and an outer diameter relative to the axis, the outer diameter has a slope angle that is less than 10 degrees relative to the axis, and wherein the inner diameter of the core flowpath increases through the low pressure compressor along a fluid flow direction; and
wherein the fan delivers air into a bypass duct, and a portion of air into the compressor section, with a bypass ratio defined as the volume of air delivered into the bypass duct compared to the volume of air delivered into the compressor section, and the bypass ratio being greater than 10.

15. The gas turbine engine of claim 14 , wherein the slope angle is between 5 degrees and 7 degrees.

16. The gas turbine engine of claim 15 , wherein the slope angle slopes toward the axis along a fluid flow direction of the core flowpath.

17. The gas turbine engine of claim 16 , wherein the outer diameter is defined by one of a plurality of rotor blades of the low pressure compressor.

18. The gas turbine engine of claim 15 , wherein a pressure ratio across the fan section is less than 1.6 across the fan blade alone.

19. The gas turbine engine of claim 18 , wherein the low pressure turbine includes an inlet, an outlet, a pressure ratio greater than 5, the pressure ratio of the low pressure turbine being pressure measured prior to the inlet as related to pressure at the outlet prior to an exhaust nozzle.

20. The gas turbine engine of claim 19 , wherein:
the geared architecture is an epicyclic gear train;
a gear reduction ratio of the epicyclic gear train is greater than 2.25;
the gas turbine engine is a two-spool turbofan;
the slope angle slopes toward the axis along a fluid flow direction of the core flowpath; and
the fan section has only a single fan stage comprising the fan.

Assignees

• Raytheon Tech Corp

Inventors

• Brilliant Lisa I
• Rose Becky E
• Dong Yuan
• Balamucki Stanley J

Classifications

• F02K3/06

  with front fan · CPC title

• F02K3/04

  the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type · CPC title

• F04D19/028

  Layout of fluid flow through the stages · CPC title

• F04D29/547 Primary

  having a special shape in order to influence fluid flow · CPC title

• F05D2260/606

  Bypassing the fluid · CPC title