US-12618378-B2 - Gas turbine engine

US12618378B2US 12618378 B2US12618378 B2US 12618378B2US-12618378-B2

Abstract

A gas turbine engine includes a turbomachine having a compressor section, a combustion section, and a turbine section, the compressor section having a high-pressure compressor defining a high-pressure compressor exit area (A HPCExit ) in square inches. The high-pressure compressor includes a high-pressure compressor flowpath and a plurality of stages. A bleed system includes a plurality of bleed flowpaths that direct compressed air from the high-pressure compressor flowpath. At least two of the bleed flowpaths are at successive stages of the plurality of stages. The gas turbine engine defines a redline exhaust gas temperature (EGT) in degrees Celsius, a total sea level static thrust output (Fn Total ) in pounds, and a corrected specific thrust, wherein the corrected specific thrust is greater than or equal to 42 and less than or equal to 90, the corrected specific determined as follows: Fn Total ×EGT/(A HPCExit 2 ×1000).

Inventors

Daniel Alan NIERGARTH
Jeffrey Donald Clements
Jeffrey S. Spruill
Erich Alois Krammer
Matthew Kenneth MacDonald
Scott Alan Schimmels

Assignees

GENERAL ELECTRIC COMPANY

Dates

Publication Date: 20260505
Application Date: 20250121

Claims (20)

1 . A gas turbine engine comprising: a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the compressor section having a high-pressure compressor defining a high-pressure compressor exit area (A HPCExit ) in square inches, the high-pressure compressor including a high-pressure compressor flowpath and a plurality of stages of high-pressure compressor rotor blades and high-pressure compressor stator vanes; and a bleed system comprising a plurality of bleed flowpaths including at least three bleed flowpaths in fluid communication with the high-pressure compressor flowpath, the plurality of bleed flowpaths directing compressed air from the high-pressure compressor flowpath, wherein at least two of the bleed flowpaths are at successive stages of the plurality of stages, wherein the gas turbine engine defines a redline exhaust gas temperature (EGT) in degrees Celsius, a total sea level static thrust output (Fn Total ) in pounds, and a corrected specific thrust, wherein the corrected specific thrust is greater than or equal to 42 and less than or equal to 90, the corrected specific thrust determined as follows: Fn Total ×EGT/(A HPCExit 2 ×1000).
2 . The gas turbine engine of claim 1 , wherein the plurality of bleed flowpaths include a first bleed flowpath at a third stage of the plurality of stages, a second bleed flowpath at a fourth stage of the plurality of stages, and a third bleed flowpath at a sixth stage of the plurality of stages.
3 . The gas turbine engine of claim 1 , wherein the plurality of bleed flowpaths include a first bleed flowpath directed to one or more aircraft systems, a second bleed flowpath directed to a low-pressure turbine of the gas turbine engine, and a third bleed flowpath directed to a high-pressure turbine of the gas turbine engine.
4 . The gas turbine engine of claim 3 , wherein the plurality of bleed flowpaths include a fourth bleed flowpath at a high-pressure compressor diffuser of the high-pressure compressor directed to at least one of the one or more aircraft systems or the low-pressure turbine.
5 . The gas turbine engine of claim 1 , wherein the high-pressure compressor further comprises an outer high-pressure compressor casing and an inner high-pressure compressor casing, and the bleed system further comprises one or more bleed plenums defined between the outer high-pressure compressor casing and the inner high-pressure compressor casing, the one or more bleed plenums defining a portion of the plurality of bleed flowpaths.
6 . The gas turbine engine of claim 5 , wherein the one or more bleed plenums include a first bleed plenum defining a first bleed flowpath, a second bleed plenum defining a second bleed flowpath, and a third bleed plenum defining a third bleed flowpath.
7 . The gas turbine engine of claim 6 , further comprising a flexible seal that separates the first bleed plenum and the second bleed plenum such that the first bleed flowpath is fluidly separate from the second bleed flowpath.
8 . The gas turbine engine of claim 6 , further comprising a rigid support member that separates the second bleed plenum and the third bleed plenum such that the second bleed flowpath is fluidly separate from the third bleed flowpath.
9 . The gas turbine engine of claim 6 , further comprising a plurality of bleed ports disposed through the inner high-pressure compressor casing and a plurality of bleed outlets disposed through the outer high-pressure compressor casing, the plurality of bleed ports and the plurality of bleed outlets being in fluid communication with the one or more bleed plenums.
10 . The gas turbine engine of claim 9 , wherein the plurality of bleed ports and the plurality of bleed outlets are sized and located in the high-pressure compressor such that the first bleed flowpath recovers more than 10% of a dynamic pressure of the first bleed flowpath into a static pressure within the first bleed flowpath.
11 . The gas turbine engine of claim 10 , wherein the plurality of bleed ports and the plurality of bleed outlets are sized and located in the high-pressure compressor such that a pressure ratio of a total pressure of the second bleed flowpath (P second_bleed ) to a total pressure of the first bleed flowpath (P first_bleed ) is less than a pressure ratio of a stage at which the second bleed flowpath is located.
12 . The gas turbine engine of claim 10 , wherein the plurality of bleed ports and the plurality of bleed outlets are sized and located in the high-pressure compressor such that the second bleed flowpath recovers less than 10% of a dynamic pressure of the second bleed flowpath into a static pressure of the second bleed flowpath.
13 . A method of operating a gas turbine engine, the method comprising: operating the gas turbine engine at a takeoff power level, the gas turbine engine having a turbomachine with a high-pressure compressor having a plurality of stages and defining a high-pressure compressor exit area (A HPCExit ) in square inches, the high-pressure compressor including a plurality of bleed flowpaths including at least three bleed flowpaths in fluid communication with a high-pressure compressor flowpath and at least two of the bleed flowpaths are at successive stages of the plurality of stages, the gas turbine engine defining a redline exhaust gas temperature (EGT) in degrees Celsius, a total sea level static thrust output (Fn Total ) in pounds, and a corrected specific thrust; directing compressed air through the high-pressure compressor flowpath of the high-pressure compressor; directing a first portion of the compressed air through a first bleed flowpath of the plurality of bleed flowpaths; directing a second portion of the compressed air through a second bleed flowpath of the plurality of bleed flowpaths; and directing a third portion of the compressed air through a third bleed flowpath of the plurality of bleed flowpaths, wherein the corrected specific thrust is greater than or equal to 42 and less than or equal to 90, the corrected specific thrust determined as follows: Fn Total ×EGT/(A HPCExit 2 ×1000).
14 . The method of claim 13 , wherein the gas turbine engine includes an inner high-pressure compressor casing and an outer high-pressure compressor casing, and the method further comprises: directing the first portion of the compressed air through a first bleed port in the inner high-pressure compressor casing of the high-pressure compressor, through a first bleed plenum defined between the inner high-pressure compressor casing and the outer high-pressure compressor casing, and through a first bleed outlet in the outer high-pressure compressor casing of the high-pressure compressor; directing the second portion of the compressed air through a second bleed port in the inner high-pressure compressor casing, through a second bleed plenum defined between the inner high-pressure compressor casing and the outer high-pressure compressor casing, and through a second bleed outlet in the outer high-pressure compressor casing; and directing the third portion of the compressed air through a third bleed port in the inner high-pressure compressor casing, through a third bleed plenum defined between the inner high-pressure compressor casing and the outer high-pressure compressor casing, and through a third bleed outlet in the outer high-pressure compressor casing.
15 . The method of claim 13 , wherein the gas turbine engine includes a third stage of the plurality of stages, a fourth stage of the plurality of stages, and a sixth stage of the plurality of stages, and the method further comprises directing the first portion of the compressed air from the third stage of the plurality of stages through the first bleed flowpath, directing the second portion of the compressed air from the fourth stage of the plurality of stages through the second bleed flowpath, and directing the third portion of the compressed air from the sixth stage of the plurality of stages through the third bleed flowpath.
16 . The method of claim 13 , wherein the first bleed flowpath recovers more than 10% of a dynamic pressure of the first bleed flowpath into a static pressure within the first bleed flowpath.
17 . The method of claim 13 , wherein a pressure ratio of a total pressure of the second bleed flowpath (P second_bleed ) to a total pressure of the first bleed flowpath (P first_bleed ) is less than a pressure ratio of a stage at which the second bleed flowpath is located.
18 . The method of claim 13 , wherein the second bleed flowpath recovers less than 10% of a dynamic pressure of the second bleed flowpath into a static pressure of the second bleed flowpath.
19 . The method of claim 13 , further comprising directing the first portion of the compressed air to one or more aircraft systems, directing the second portion of the compressed air to a low-pressure turbine of the gas turbine engine, and directing the third portion of the compressed air to a high-pressure turbine of the gas turbine engine.
20 . The method of claim 19 , wherein the high-pressure compressor includes a high-pressure compressor diffuser, and the method further comprises directing a fourth portion of the compressed air from the high-pressure compressor diffuser of the high-pressure compressor, and directing the fourth portion of the compressed air to at least one of the one or more aircraft systems or the low-pressure turbine.

Description

CROSS-REFERENCE TO RELATED APPLICATIONS This application is a continuation-in-part patent application of U.S. patent application Ser. No. 18/481,515 filed Oct. 5, 2023, which is a continuation-in-part patent application of U.S. patent application Ser. No. 17/978,629 filed Nov. 1, 2022. Each of these applications are hereby incorporated by reference in their entirety. FIELD The present disclosure relates to a gas turbine engine. BACKGROUND A gas turbine engine typically includes a fan and a turbomachine. The turbomachine generally includes an inlet, one or more compressors, a combustor, and at least one turbine. The compressors compress air which is channeled to the combustor where it is mixed with fuel. The mixture is then ignited for generating hot combustion gases. The combustion gases are channeled to the turbine(s) which extracts energy from the combustion gases for powering the compressor(s), as well as for producing useful work to propel an aircraft in flight. The turbomachine is mechanically coupled to the fan for driving the fan during operation. 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 schematic, cross-sectional view of a three-stream engine in accordance with an exemplary embodiment of the present disclosure. FIG. 2 is a close-up, schematic view of the exemplary three-stream engine of FIG. 1 with a cooled cooling air system in accordance with an exemplary embodiment of the present disclosure. FIG. 3 is a close-up view of an aft-most stage of high-pressure compressor rotor blades within the exemplary three-stream engine of FIG. 1. FIG. 4 is a close-up, schematic view of the exemplary three-stream engine of FIG. 1 showing the cooled cooling air system of FIG. 2. FIG. 5 is a schematic view of a thermal transport bus of the present disclosure. FIG. 6 is a table depicting numerical values showing the relationships between various parameters in accordance with various example embodiments of the present disclosure. FIG. 7 is a graph depicting a range of corrected specific thrust values and redline exhaust gas temperature values of gas turbine engines in accordance with various example embodiments of the present disclosure. FIG. 8 is a schematic view of a ducted turbofan engine in accordance with an exemplary aspect of the present disclosure. FIG. 9 is a schematic, close-up view of a gas turbine engine having a cooled cooling air system in accordance with another exemplary aspect of the present disclosure. FIG. 10 is a schematic, close-up view of a gas turbine engine having a cooled cooling air system in accordance with yet another exemplary aspect of the present disclosure. FIG. 11 is a schematic, close-up view of a gas turbine engine having a cooled cooling air system in accordance with still another exemplary aspect of the present disclosure. FIG. 12 is a schematic view of a turbofan engine in accordance with another exemplary aspect of the present disclosure. FIG. 13 is a schematic cross-sectional diagram of a ducted turbofan engine, taken along a longitudinal centerline axis of the engine, according to the present disclosure. FIG. 14 is a schematic, cross-sectional view of a high-pressure compressor of the engine of FIG. 13, taken at detail 14-14 in FIG. 13, according to the present disclosure. FIG. 15 is a flow diagram of a method of operating the engine of FIGS. 13 and 14, according to the present disclosure. FIG. 16 is a flow diagram of a method of operating the engine of FIGS. 13 and 14, according to another embodiment. FIG. 17 represents, in graph form, a mass flow ratio of a first mass flow Wfirst_bleed through a first bleed flowpath to a second mass flow Wsecond_bleed through a second bleed flowpath as a function of an altitude of the engine of FIG. 13, according to 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 term “cooled cooling air system” is used herein to mean a system configured to provide a cooling airflow to one or more components exposed to a working gas flowpath of a turbomachine of a gas turbine engine at a location downstream of a combustor of the turbomachine and upstream of an exhaust nozzle of the turbomachine, the cooling airflow being in thermal communication with a heat exchanger for reducing a temperature of the cooling airflow at a location upstream of the one or more components. The cooled cooling air systems contemplated by the present disclosure may include a