Gas turbine engine with heat exchanger assembly

Patent metadata

Field	Value
Publication number	US-2025305454-A1
Application number	US-202418619703-A
Country	US
Kind code	A1
Filing date	Mar 28, 2024
Priority date	Mar 28, 2024
Publication date	Oct 2, 2025
Grant date	—

Abstract

A gas turbine engine includes a fan assembly, a turbomachine defining in part a working gas flowpath, the gas turbine engine defining a bypass passage over the turbomachine, a core cowl, and a heat exchanger assembly including a heat exchanger and a heat exchanger cowl defining a cooling air flowpath extending between a flowpath inlet in airflow communication with the bypass passage to receive a cooling airflow from the bypass passage and a flowpath outlet in airflow communication with the bypass passage to exhaust the cooling airflow back to the bypass passage, the heat exchanger positioned within the cooling air flowpath, the cooling air flowpath comprising a diffusion section located between the flowpath inlet and the heat exchanger.

First claim

What is claimed is:

1 . A gas turbine engine defining a radial direction and an axial direction, the gas turbine engine comprising:
a fan assembly comprising a fan; a turbomachine drivingly coupled to the fan and comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order and defining in part a working gas flowpath, the gas turbine engine defining a bypass passage over the turbomachine; a core cowl; and a heat exchanger assembly disposed aft of the fan in the axial direction and forward of the core cowl in the axial direction, the heat exchanger assembly comprising:
a heat exchanger; and
a heat exchanger cowl defining a cooling air flowpath extending between a flowpath inlet in airflow communication with the bypass passage to receive a cooling airflow from the bypass passage and a flowpath outlet in airflow communication with the bypass passage to exhaust the cooling airflow back to the bypass passage, the heat exchanger positioned within the cooling air flowpath, the cooling air flowpath comprising a diffusion section located between the flowpath inlet and the heat exchanger.

2 . The gas turbine engine of claim 1 , wherein the cooling air flowpath further defines an exhaust section located between the heat exchanger and the flowpath outlet.

3 . The gas turbine engine of claim 2 , wherein the turbomachine comprises an ejector positioned in airflow communication with the exhaust section, wherein the ejector is in airflow communication with a high pressure air source.

4 . The gas turbine engine of claim 3 , wherein the compressor section comprises a low pressure compressor and a high pressure compressor, wherein the turbomachine further comprises a variable bleed assembly disposed downstream of the low pressure compressor and upstream of the high pressure compressor, wherein the high pressure air source is the variable bleed assembly.

5 . The gas turbine engine of claim 1 , wherein the compressor section of the turbomachine comprises a booster, wherein the heat exchanger is disposed outward of the booster in the radial direction.

6 . The gas turbine engine of claim 5 , further comprising a booster cowl, wherein the booster cowl is disposed between the booster and the heat exchanger assembly in the radial direction.

7 . The gas turbine engine of claim 1 , wherein the fan further comprises an outlet guide vane, and wherein the heat exchanger assembly is disposed aft of the outlet guide vane in the axial direction.

8 . The gas turbine engine of claim 1 , wherein the turbomachine comprises a circumferential groove, wherein the core cowl comprises a first member, a second member, and a hinged connection therebetween, wherein the first member and the second member are coupled to the circumferential groove, and wherein the heat exchanger assembly is disposed forward of the circumferential groove in the axial direction.

9 . The gas turbine engine of claim 1 , wherein the heat exchanger is a first heat exchanger of a plurality of heat exchangers of the heat exchanger assembly.

10 . The gas turbine engine of claim 9 , wherein the heat exchanger cowl is a first heat exchanger cowl of a plurality of heat exchanger cowls of the heat exchanger assembly, wherein the cooling air flowpath is a first cooling air flowpath of a plurality of cooling air flowpaths of the heat exchanger assembly, and wherein each cooling air flowpath of the plurality of cooling air flowpaths is defined by a respective one of the plurality of heat exchanger cowls.

11 . The gas turbine engine of claim 10 , further defining a circumferential direction, and wherein the plurality of heat exchanger cowls and the plurality of heat exchangers are arranged in the circumferential direction.

12 . The gas turbine engine of claim 9 , further comprising a heat source in thermal communication with the plurality of heat exchangers.

13 . The gas turbine engine of claim 9 , wherein the plurality of heat exchangers includes at least two heat exchangers and less than or equal to 20 heat exchangers.

14 . The gas turbine engine of claim 1 , wherein the heat exchanger is one of an air-cooled oil cooler, a variable frequency generator air-cooled oil cooler, or air-air heat exchanger, or an air-cooled fluid cooler.

15 . The gas turbine engine of claim 1 , wherein the heat exchanger assembly further comprises a heat source in thermal communication with the heat exchanger.

16 . The gas turbine engine of claim 1 , wherein the diffusion section is configured to reduce a speed of air flowing to the flowpath inlet.

17 . The gas turbine engine of claim 1 , wherein the cooling air flowpath further defines an exhaust section located between the heat exchanger and the flowpath outlet, wherein the exhaust section is configured to accelerate the cooling airflow from the heat exchanger provided to the bypass passage.

18 . The gas turbine engine of claim 1 , wherein the fan is a single stage fan.

19 . The gas turbine engine of claim 1 , wherein the turbomachine defines an inlet to the working gas flowpath and comprises a circumferential groove for mounting at least a portion of the core cowl, wherein the heat exchanger cowl defines a first length along the axial direction from the flowpath inlet to the flowpath outlet, wherein the turbomachine defines a second length along the axial direction from the inlet to the circumferential groove, and wherein L 1 is greater than or equal to 10% of L 2 and is less than or equal to 60% of L 2 .

20 . The gas turbine engine of claim 19 , wherein L 1 is greater than or equal to 20% of L 2 and is less than or equal to 50% of L 2 .

Assignees

• Gen Electric

Inventors

• Miller Brandon Wayne
• Imgrund Jessica Breanna
• Barre Eric
• Gallier Kirk Douglas
• Holland John R
• Johnson Steven Douglas

Classifications

• F02K3/115

  by means of indirect heat exchange · CPC title

• F02K3/06

  with front fan · CPC title

• F02C7/14

  of fluids in the plant {, e.g. lubricant or fuel (F02C7/185 takes precedence)} · CPC title

• F02C7/141 Primary

  of working fluid · CPC title

• F02C7/18

  the medium being gaseous, e.g. air {(F02C7/125 takes precedence)} · CPC title