Turboshaft engine

Patent metadata

Field	Value
Publication number	US-12331702-B2
Application number	US-202418784568-A
Country	US
Kind code	B2
Filing date	Jul 25, 2024
Priority date	Jun 4, 2021
Publication date	Jun 17, 2025
Grant date	Jun 17, 2025

Abstract

A turboshaft engine includes a core engine, including a fan section, a compressor section, a primary combustor and a turbine section positioned within a core flow path of the gas turbine engine; a bypass splitter positioned radially outward of the core engine and configured to house the compressor section, the primary combustor and the turbine section; a bypass duct positioned radially outward of the bypass splitter; and a power spool operably coupled to the core engine and configured rotationally drive a fan included within the fan section.

First claim

What is claimed:

1. A gas turbine engine, comprising:
a core engine, including a compressor section, a primary combustor and a turbine section positioned within a core flow path of the gas turbine engine;
a bypass splitter positioned radially outward of the core engine and configured to house the compressor section, the primary combustor and the turbine section;
a bypass duct positioned radially outward of the bypass splitter;
a power spool operably coupled to the core engine;
a fan section, wherein the power spool is operably connected to a fan within the fan section and configured to rotationally drive the fan;
a supplemental combustor, wherein the supplemental combustor is a rotating detonation combustor that includes a fuel-air mixer configured to receive a compressed air and a fuel, wherein the rotating detonation combustor includes an annular structure positioned downstream of the fuel-air mixer and configured to combust the compressed air and the fuel, wherein the compressed air is provided to the rotating detonation combustor via the bypass duct, and wherein the rotating detonation combustor is positioned within an exit nozzle of the gas turbine engine, the exit nozzle positioned downstream of the bypass duct and wherein the rotating detonation combustor is configured to combust the compressed air, the fuel and an exhaust stream exiting a power turbine;
a bypass flow variable strut configured to manage or regulate a bypass flow through the bypass duct; and
a core flow variable strut configured to at least one of manage or regulate a core flow into a combustion section, wherein both the bypass flow variable strut and the core flow variable strut are positioned aft of the fan and forward of the bypass splitter.

2. The gas turbine engine of claim 1 , further comprising a high-speed spool and wherein the compressor section includes a high-pressure compressor and the turbine section includes a high-pressure turbine, the high-pressure compressor and the high-pressure turbine being interconnected via the high-speed spool.

3. The gas turbine engine of claim 2 , further comprising a low-speed spool and wherein the compressor section includes a low-pressure compressor and the turbine section includes a low-pressure turbine, the low-pressure compressor and the low-pressure turbine being interconnected via the low-speed spool.

4. The gas turbine engine of claim 3 , wherein the power spool includes the power turbine positioned downstream of the low-pressure turbine.

5. The gas turbine engine of claim 4 , wherein the power spool includes a power shaft configured to interconnect a fan gearbox to the power turbine, the fan gearbox operably connected to the fan.

6. A turboshaft engine for an aircraft, comprising:
a core engine, including a fan section, a compressor section, a primary combustor and a turbine section positioned within a core flow path of the turboshaft engine;
a bypass splitter positioned radially outward of the core engine and configured to house the compressor section, the primary combustor and the turbine section;
a bypass duct positioned radially outward of the bypass splitter;
a power spool operably coupled to the core engine and configured rotationally drive a fan included within the fan section;
a main spool, wherein the compressor section includes a compressor and the turbine section includes a primary turbine, the compressor and the primary turbine being interconnected via the main spool, wherein the power spool includes a power turbine positioned downstream of the primary turbine, and wherein the power spool includes a power shaft configured to interconnect a fan gearbox to the power turbine, the fan gearbox operably connected to the fan;
a rotating detonation combustor configured to receive a compressed air and a fuel, wherein the compressed air is provided to the rotating detonation combustor via the bypass duct and wherein the rotating detonation combustor is positioned within an exit nozzle of the turboshaft engine, the exit nozzle positioned downstream of the bypass duct and wherein the rotating detonation combustor is configured to combust the compressed air, the fuel and an exhaust stream exiting the power turbine;
a bypass flow variable strut configured to manage or regulate a bypass flow through the bypass duct; and
a core flow variable strut configured to at least one of manage or regulate a core flow into a combustion section, wherein both the bypass flow variable strut and the core flow variable strut are positioned aft of the fan and forward of the bypass splitter.

Assignees

• Rtx Corp

Inventors

• Burd Steven W

Classifications

• F02C9/18

  by bleeding, bypassing or acting on variable working fluid interconnections between turbines or compressors or their stages {(F02C3/113 takes precedence)} · CPC title

• F05D2240/35

  Combustors or associated equipment · CPC title

• F05D2220/32

  in gas turbines · CPC title

• F05D2260/606

  Bypassing the fluid · CPC title

• F02C7/36

  Power transmission arrangements between the different shafts of the gas turbine plant, or between the gas-turbine plant and the power user ({F02C3/107 - F02C3/13 and} F02C7/32 take precedence; couplings for transmitting rotation F16D; gearing in general F16H) · CPC title