US-12618368-B2 - Gas turbine engine

US12618368B2US 12618368 B2US12618368 B2US 12618368B2US-12618368-B2

Abstract

A gas turbine engine is provided. The gas turbine engine includes: a fan; a turbomachine drivingly coupled to the fan and defining in part a working gas flowpath, the gas turbine engine defining a bypass passage over the turbomachine, the turbomachine defining an annular cooling passage extending between a CP inlet and a CP outlet, the CP inlet in airflow communication with the working gas flowpath and the CP outlet in airflow communication with the bypass passage; and a variable bleed assembly including a variable bleed duct extending between a VB inlet and a VB outlet, the VB inlet in airflow communication with the working gas flowpath at a location downstream of the CP inlet and the VB outlet in airflow communication with the annular cooling passage for urging an airflow through the cooling passage.

Inventors

Brandon Wayne Miller
Andrew Hudecki
Eric Barre

Assignees

GENERAL ELECTRIC COMPANY

Dates

Publication Date: 20260505
Application Date: 20221130

Claims (19)

1 . A gas turbine engine comprising: a fan assembly comprising a fan at a forward end of the gas turbine engine, the fan being a first stage of compression for the gas turbine engine; 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, the turbomachine defining an annular cooling passage extending between a CP inlet and a CP outlet, the CP inlet in airflow communication with the working gas flowpath and the CP outlet in airflow communication with the bypass passage, wherein the turbomachine defines an inlet to the working gas flowpath located downstream of the fan, wherein the compressor section defines a second stage of compression for the gas turbine engine, wherein the CP inlet is in airflow communication with the working gas flowpath at a location upstream of the second stage of compression, and wherein no intermediate stages of compression are located between the first stage of compression and the second stage of compression; and a variable bleed assembly comprising a variable bleed duct extending between a VB inlet and both a VB outlet and supplementary VB outlet, the VB inlet in airflow communication with the working gas flowpath at a location downstream of the CP inlet, the VB outlet in airflow communication with the annular cooling passage for urging an airflow through the annular cooling passage and the supplementary VB outlet is in direct airflow communication with the bypass passage and is separated and downstream from the CP outlet.
2 . The gas turbine engine of claim 1 , wherein the compressor section comprises a compressor, and wherein the VB inlet is in airflow communication with the working gas flowpath at a location downstream of the compressor.
3 . The gas turbine engine of claim 2 , wherein the compressor is a low pressure compressor.
4 . The gas turbine engine of claim 3 , wherein the compressor section further comprises a high pressure compressor, wherein the location is upstream of the high pressure compressor.
5 . The gas turbine engine of claim 1 , wherein the variable bleed assembly comprises a variable bleed valve for varying an amount of a bleed airflow through the variable bleed duct.
6 . The gas turbine engine of claim 5 , further comprising: a controller operably coupled to the variable bleed valve, wherein the controller is configured to actuate the variable bleed assembly to increase the amount of the bleed airflow through the variable bleed duct in response to an operating condition of the gas turbine engine to increase an amount of the airflow through the annular cooling passage.
7 . The gas turbine engine of claim 1 , wherein the VB outlet forms at least in part an ejector.
8 . The gas turbine engine of claim 1 , wherein substantially all of the airflow through the variable bleed duct is provided through the VB outlet to the annular cooling passage.
9 . The gas turbine engine of claim 1 , wherein the turbomachine comprises a heat exchanger in thermal communication with the airflow through the annular cooling passage.
10 . The gas turbine engine of claim 9 , wherein the VB outlet is in airflow communication with the annular cooling passage at a location downstream of the heat exchanger.
11 . The gas turbine engine of claim 1 , wherein the fan of the fan assembly is a single stage fan.
12 . The gas turbine engine of claim 1 , wherein the fan of the fan assembly is an open rotor fan.
13 . The gas turbine engine of claim 1 , wherein the compressor section comprises a low pressure compressor and a high pressure compressor.
14 . A method of operating a gas turbine engine comprising a fan assembly and a turbomachine drivingly coupled to a fan of the fan assembly, the fan at a forward end of the gas turbine engine and being a first stage of compression for the gas turbine engine, the method comprising: receiving data indicative of an operating condition of the gas turbine engine; varying an amount of a variable bleed airflow through a variable bleed duct, the variable bleed duct extending between a VB inlet and both a VB outlet and a supplementary VB outlet where the VB inlet is in airflow communication with a working gas flowpath at a location downstream of a CP inlet; and providing to an annular cooling passage, via the VB outlet, the amount of the variable bleed airflow in response to the data, the annular cooling passage extending between the CP inlet in airflow communication with the working gas flowpath of the turbomachine and a CP outlet in airflow communication with a bypass passage of the gas turbine engine, wherein the turbomachine defines an inlet to the working gas flowpath located downstream of the fan, wherein a compressor section defines a first second stage of compression downstream of the inlet, wherein the CP inlet in airflow communication with the working gas flowpath at a location upstream of the first second stage of compression, wherein no intermediate stages of compression are located between the first stage of compression and the second stage of compression, and wherein the supplementary VB outlet is in direct airflow communication with the bypass passage and is separated and downstream from the CP outlet.
15 . The method of claim 14 , wherein the operating condition is a low fan power operating condition, and wherein varying the amount of variable bleed airflow through the variable bleed duct provided to the annular cooling passage comprises increasing the amount of variable bleed airflow through the variable bleed duct provided to the annular cooling passage.
16 . The method of claim 15 , wherein the low fan power operating condition is a ground idle operating condition or a flight idle descent operating condition.
17 . The method of claim 14 , wherein the operating condition is indicative of an ambient temperature.
18 . The method of claim 14 , wherein a variable bleed assembly comprises a variable bleed valve for varying the amount of the variable bleed airflow through the variable bleed duct.
19 . The method of claim 14 , wherein the turbomachine comprises a heat exchanger in thermal communication with an airflow through the annular cooling passage.

Description

CROSS-REFERENCE TO RELATED APPLICATIONS This application is a non-provisional application claiming the benefit of priority under 35 U.S.C. § 119(e) to U.S. Provisional Application No. 63/411,180, filed Sep. 29, 2022, which is hereby incorporated by reference in its entirety. FIELD The present disclosure relates to a gas turbine engine. BACKGROUND A gas turbine engine generally includes a turbomachine and a rotor assembly. Gas turbine engines, such as turbofan engines, may be used for aircraft propulsion. In the case of a turbofan engine, the rotor assembly may be configured as a fan assembly, and the fan assembly may be enclosed by an outer nacelle. The outer nacelle may define a bypass passage with the turbomachine. Generally, improvements to a turbofan engine in the fields of thermal management and aerodynamics would be welcomed in the art. BRIEF DESCRIPTION OF THE DRAWINGS A full and enabling disclosure of the present disclosure, including the best mode thereof, directed to one of ordinary skill in the art, is set forth in the specification, which makes reference to the appended figures, in which: FIG. 1 is a cross-sectional view of a gas turbine engine in accordance with an exemplary aspect of the present disclosure. FIG. 2 is a schematic view of a portion of a turbomachine in accordance with an exemplary embodiment of the present disclosure. FIG. 3 is a schematic, cross-sectional view of a cooling passage of the turbomachine of FIG. 2 in accordance with an exemplary embodiment of the present disclosure. FIG. 4 is a schematic view of a portion of a turbomachine in accordance with another exemplary embodiment of the present disclosure. FIG. 5 is a schematic view of a hood of the exemplary turbomachine of FIG. 4. FIG. 6 is a schematic view of a hood of a turbomachine in accordance with an exemplary embodiment of the present disclosure. FIG. 7 is a schematic view of a portion of a turbomachine in accordance with another exemplary embodiment of the present disclosure. FIG. 8 is a schematic view of an ejector in accordance with another exemplary embodiment of the present disclosure. FIG. 9 is a schematic, cross-sectional view of a cooling passage and variable bleed assembly in accordance with an exemplary embodiment of the present disclosure. FIG. 10 is a flow diagram of a method of operating a gas turbine engine in accordance with an exemplary aspect of the present disclosure. FIG. 11 is a cross-sectional view of a gas turbine engine in accordance with another exemplary aspect of the present disclosure. FIG. 12 is a schematic view of a portion of a turbomachine in accordance with another exemplary embodiment of the present disclosure. FIG. 13 is an isometric, cross-sectional view of a variable bleed assembly in accordance with an exemplary embodiment of the present disclosure. FIG. 14 is a schematic view of a hinged cowl assembly. FIG. 15 is a schematic view of a portion of a turbomachine in accordance with another exemplary embodiment of the present disclosure. DETAILED DESCRIPTION Reference will now be made in detail to present embodiments of the disclosure, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerical and letter designations to refer to features in the drawings. Like or similar designations in the drawings and description have been used to refer to like or similar parts of the disclosure. The word “exemplary” is used herein to mean “serving as an example, instance, or illustration.” Any implementation described herein as “exemplary” is not necessarily to be construed as preferred or advantageous over other implementations. Additionally, unless specifically identified otherwise, all embodiments described herein should be considered exemplary. The singular forms “a”, “an”, and “the” include plural references unless the context clearly dictates otherwise. The term “at least one of” in the context of, e.g., “at least one of A, B, and C” refers to only A, only B, only C, or any combination of A, B, and C. The term “turbomachine” refers to a machine including one or more compressors, a heat generating section (e.g., a combustion section), and one or more turbines that together generate a torque output. The term “gas turbine engine” refers to an engine having a turbomachine as all or a portion of its power source. Example gas turbine engines include turbofan engines, turboprop engines, turbojet engines, turboshaft engines, etc., as well as hybrid-electric versions of one or more of these engines. The term “combustion section” refers to any heat addition system for a turbomachine. For example, the term combustion section may refer to a section including one or more of a deflagrative combustion assembly, a rotating detonation combustion assembly, a pulse detonation combustion assembly, or other appropriate heat addition assembly. In certain example embodiments, the combustion section may include an annular combustor, a can combustor, a cannula