US-12618384-B2 - Gas turbine engine having a bypass passage, a fan duct, and a core duct

US12618384B2US 12618384 B2US12618384 B2US 12618384B2US-12618384-B2

Abstract

A gas turbine engine includes: a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the turbomachine defining an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct; a primary fan driven by the turbomachine; and a secondary fan located downstream of the primary fan within the inlet duct, the gas turbine engine defining a thrust to power airflow ratio between 3.5 and 100 and a core bypass ratio between 0.1 and 10, wherein the thrust to power airflow ratio is a ratio of an airflow through a bypass passage over the turbomachine plus an airflow through the fan duct to an airflow through the core duct, wherein the core bypass ratio is a ratio of the airflow through the fan duct to the airflow through the core duct.

Inventors

Brandon Wayne Miller
Randy M. Vondrell
David Marion Ostdiek
Craig William Higgins
Alexander Kimberley Simpson
Syed Arif Khalid

Assignees

GENERAL ELECTRIC COMPANY

Dates

Publication Date: 20260505
Application Date: 20250110

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 turbomachine defining an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct; a primary fan driven by the turbomachine; and a secondary fan located downstream of the primary fan within the inlet duct, the gas turbine engine defining a thrust to power airflow ratio between 3.5 and 100 and a core bypass ratio between 0.1 and 10, wherein the thrust to power airflow ratio is a ratio of an airflow through a bypass passage over the turbomachine plus an airflow through the fan duct to an airflow through the core duct, and wherein the core bypass ratio is a ratio of the airflow through the fan duct to the airflow through the core duct, and wherein the primary fan includes: a plurality of blades, each of the plurality of blades comprising: a blade root proximal to an axis of rotation; a blade tip remote from the axis; and a blade span measured between the blade root and the blade tip, wherein a circumferential averaged tangential velocity, Cu, and a radius of airflow, R, and a loading of a respective blade of the plurality of blades at R is defined as a change in RCu, ΔRCu, over a row of blades in the plurality of blades, and wherein the primary fan has a load distribution such that, at any location between the blade root of a respective blade and 30% span, a value of ΔRCu is greater than or equal to 60% of a peak ΔRCu.
2 . The gas turbine engine of claim 1 , wherein the thrust to power airflow ratio and the core bypass ratio are defined when the gas turbine engine is operated at a rated speed during standard day operating conditions.
3 . The gas turbine engine of claim 1 , wherein the secondary fan is a single stage secondary fan.
4 . The gas turbine engine of claim 1 , wherein the secondary fan is a multi-stage secondary fan.
5 . The gas turbine engine of claim 1 , wherein the primary fan, the secondary fan, the compressor section, the combustion section, and the turbine section are arranged in serial flow order.
6 . The gas turbine engine of claim 1 , wherein the primary fan is at least partially covered by a shroud.
7 . The gas turbine engine of claim 1 , further including a plurality of vanes, each vane of the plurality of vanes comprising a vane root proximal to the axis, a vane tip remote from the axis, and a vane span measured between the vane root and the vane tip configured to impart a change in tangential velocity of the air opposite to that imparted by the primary fan.
8 . The gas turbine engine of claim 7 , wherein at least one of the primary fan or the plurality of vanes is unducted.
9 . The gas turbine engine of claim 7 , wherein the plurality of vanes is positioned upstream of the primary fan.
10 . The gas turbine engine of claim 7 , wherein the plurality of vanes is positioned downstream of the primary fan.
11 . The gas turbine engine of claim 7 , wherein at least one of the plurality of vanes comprises a shroud distally from the axis.
12 . The gas turbine engine of claim 7 , wherein a circumferential averaged tangential velocity of an airflow aft of the gas turbine engine is lower than a change in the circumferential averaged tangential velocity of the primary fan over a majority of the span of the vanes.
13 . The gas turbine engine of claim 1 , wherein the gas turbine engine is a propeller system.
14 . The gas turbine engine of claim 1 , wherein the gas turbine engine is an open rotor system.
15 . The gas turbine engine of claim 1 , wherein the blades in the plurality of blades have a characteristic selected from the group consisting of: i) a blade camber at 30% span is at least 90% of a maximum blade camber between 50% span and 100% span, ii) the blade camber at 0% span is at least 110% of the maximum blade camber between 50% span and 100% span, and iii) combinations thereof.
16 . The gas turbine engine of claim 1 , wherein, at 30% span, the value of ΔRCu is greater than or equal to 70% of the peak ΔRCu.
17 . A method of operating a gas turbine engine, comprising: operating the gas turbine engine at a rated speed, wherein operating the gas turbine engine at the rated speed comprises operating the gas turbine engine to define a thrust to power airflow ratio between 3.5 and 100 and a core bypass ratio between 0.1 and 5, wherein the thrust to power airflow ratio is a ratio of an airflow through a bypass passage over a turbomachine of the gas turbine engine plus an airflow through a fan duct to an airflow through a core duct, and wherein the core bypass ratio is a ratio of the airflow through the fan duct to the airflow through the core duct, and wherein the primary fan includes: a plurality of blades, each of the plurality of blades comprising: a blade root proximal to an axis of rotation; a blade tip remote from the axis; and a blade span measured between the blade root and the blade tip, wherein a circumferential averaged tangential velocity, Cu, and a radius of airflow, R, and a loading of a respective blade of the plurality of blades at R is defined as a change in RCu, ΔRCu, over a row of blades in the plurality of blades, and wherein the primary fan has a load distribution such that, at any location between the blade root of a respective blade and 30% span, a value of ΔRCu is greater than or equal to 60% of a peak ΔRCu.
18 . The method of claim 17 , wherein the gas turbine engine further includes a plurality of vanes, each vane of the plurality of vanes comprising a vane root proximal to the axis, a vane tip remote from the axis, and a vane span measured between the vane root and the vane tip configured to impart a change in tangential velocity of the air opposite to that imparted by the primary fan.
19 . The method of claim 18 , wherein a circumferential averaged tangential velocity of an airflow aft of the gas turbine engine is lower than a change in the circumferential averaged tangential velocity of the primary fan over a majority of the span of the vanes.
20 . The method of claim 17 , wherein blades of the plurality of blades have a characteristic selected from the group consisting of: i) a blade camber at 30% span is at least 90% of a maximum blade camber between 50% span and 100% span, ii) the blade camber at 0% span is at least 110% of the maximum blade camber between 50% span and 100% span, and iii) combinations thereof.

Description

CROSS-REFERENCE TO RELATED APPLICATIONS This application is a continuation in part application of U.S. application Ser. No. 18/675,270, filed May 28, 2024, which is a continuation application of U.S. application Ser. No. 17/879,384 filed Aug. 2, 2022. Each of these applications is hereby incorporated by reference in their entireties. FIELD The present disclosure relates to a gas turbine engine with a third stream. 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. FIG. 3 is a close-up view of an area surrounding a leading edge of a core cowl of the exemplary three-stream engine of FIG. 2. FIGS. 4A through 4H are tables depicting numerical values showing the relationships between various parameters in accordance with various example embodiments of the present disclosure. FIGS. 5A through 5D are graphs depicting a range of thrust to power airflow ratios and core bypass ratios in accordance with various example embodiments of the present disclosure. FIG. 6 is a schematic view of a turboprop engine in accordance with an exemplary aspect of the present disclosure. FIG. 7 is a schematic view of a direct drive, ducted, turbofan engine in accordance with an exemplary aspect of the present disclosure. FIG. 8 is a schematic view of a geared, ducted, turbofan engine in accordance with an exemplary aspect of the present disclosure. FIG. 9 is a schematic view of a geared, ducted, turbofan engine in accordance with another exemplary aspect of the present disclosure. FIG. 10 is a schematic view of an unducted gas turbine engine in accordance with another exemplary aspect of the present disclosure. FIG. 11 is a schematic view of an unducted gas turbine engine in accordance with yet another exemplary aspect of the present disclosure. FIG. 12 is a schematic view of an unducted gas turbine engine in accordance with still another exemplary aspect of the present disclosure. FIG. 13 shows an elevational cross-sectional view of an exemplary unducted thrust producing system. FIG. 14 is an illustration of an alternative embodiment of an exemplary vane assembly for an unducted thrust producing system. FIG. 15 depicts vector diagrams illustrating circumferential averaged tangential velocity through both rows for two exemplary embodiments. FIG. 16 depicts graphically the aerodynamic rotor load distribution of two exemplary embodiments of an unducted thrust producing system in comparison with a conventional configuration. FIG. 17 depicts graphically the exit swirl velocity and axial velocity for two exemplary embodiments of an unducted thrust producing system in comparison with two conventional configurations. FIG. 18 depicts graphically how various parameters such as camber and stagger angle are defined with respect to a blade or vane FIG. 19 depicts graphically representative parameters associated with an exemplary embodiment of an airfoil blade in comparison with a conventional airfoil blade. FIG. 20 is an elevational view of an exemplary airfoil blade for an unducted thrust producing system with section line locations 1 through 11 identified. FIGS. 21-31 are cross-sectional illustrations of the exemplary airfoil blade of FIG. 20 at section line locations 1 through 11 in comparison with analogous sections through the conventional airfoil blade referenced previously. 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 ot