Compound helicopters having hybrid propulsion engines

Patent metadata

Field	Value
Publication number	US-11267579-B2
Application number	US-201916365573-A
Country	US
Kind code	B2
Filing date	Mar 26, 2019
Priority date	Mar 26, 2019
Publication date	Mar 8, 2022
Grant date	Mar 8, 2022

Abstract

A hybrid propulsion engine for a rotorcraft includes a core turboshaft engine having a gas path and an output shaft that provides torque to a main rotor. A fan module is disposed relative to the core turboshaft engine and is coupled to the output shaft. The fan module has a bypass air path that is independent of the gas path. A thrust nozzle is configured to mix exhaust gases from the core turboshaft engine with bypass air from the fan module and to discharge the mixture to provide propulsive thrust. In a turboshaft configuration, the fan module is closed to prevent the flow of bypass air therethrough such that the thrust nozzle does not provide propulsive thrust. In a turboshaft and turbofan configuration, the fan module is open allowing the flow of bypass air therethrough such that the thrust nozzle provides propulsive thrust, thereby supplying propulsion compounding for the rotorcraft.

First claim

What is claimed is:

1. A compound rotorcraft comprising:
a fuselage having a tail cone;
first and second wings coupled to the fuselage and configured to provide lift compounding responsive to forward airspeed;
a first tail boom coupled to the first wing and having a first aft end;
a second tail boom coupled to the second wing and having a second aft end, the first aft end and the second aft end coupled together by an empennage, the tail cone located between the first tail boom and the second tail boom;
a tail rotor rotatably coupled to the empennage;
a main rotor coupled to and rotatable relative to the fuselage; and
a hybrid propulsion engine deposed within the fuselage and including:
a core turboshaft engine having a gas path including an inlet air stage, a compression stage, a combustion stage, a turbine stage and an exhaust stage, the core turboshaft engine including an output shaft that provides torque to the main rotor;
a fan module disposed relative the core turboshaft engine and coupled to the output shaft, the fan module having a bypass air path that is independent of the gas path, the bypass air path including a bypass air inlet stage, a bypass air compression stage and a bypass air exhaust stage, the fan module having a closed configuration to prevent bypass air from flowing therethrough and an open configuration to allow bypass air to flow therethrough; and
a thrust nozzle disposed within the tail cone in a non-parallel arrangement relative to the first tail boom and the second tail boom, the thrust nozzle configured to mix the exhaust gases from the exhaust stage with the bypass air from the bypass air exhaust stage and to discharge the exhaust gases and bypass air mixture from the tail cone with a downward angle relative the first tail boom and the second tail boom to provide propulsive thrust;
wherein, in a turboshaft configuration of the hybrid propulsion engine, the fan module is in the closed configuration such that the thrust nozzle does not provide propulsive thrust; and
wherein, in a turboshaft and turbofan configuration of the hybrid propulsion engine, the fan module is in the open configuration such that the thrust nozzle provides propulsive thrust, thereby supplying propulsion compounding for the rotorcraft.

2. The compound rotorcraft as recited in claim 1 wherein the main rotor further comprises a single main rotor.

3. The compound rotorcraft as recited in claim 1 wherein the inlet air stage is configured to draw in ambient air;
wherein the compression stage includes a compressor configured to compress the ambient air from the inlet air stage;
wherein the combustion stage is configured to mix the compressed air from the compression stage with fuel and ignite the compressed air and fuel mixture to form hot combustion gases;
wherein the turbine stage includes a turbine operated responsive to the flow of hot combustion gases from the combustion stage, the turbine driving the compressor and the output shaft; and
wherein the exhaust stage discharges exhaust gases from the turbine stage.

4. The compound rotorcraft as recited in claim 3 wherein the compressor further comprises a multi-stage axial compressor having a pressure ratio of between about 10 to 1 and about 20 to 1.

5. The compound rotorcraft as recited in claim 3 wherein the turbine further comprises a compressor turbine that drives the compressor and a power turbine that drives the output shaft and wherein the compressor turbine rotates independently of the power turbine.

6. The compound rotorcraft as recited in claim 5 wherein the compressor turbine further comprises a multi-stage compressor turbine and the power turbine further comprises a multi-stage power turbine.

7. The compound rotorcraft as recited in claim 1 wherein the fan module further comprises a duct assembly and wherein the bypass air path further comprises an annular bypass air path between the core turboshaft engine and the duct assembly.

8. The compound rotorcraft as recited in claim 1 wherein the fan module further comprises a plurality of inlet doors actuatable between open and closed positions to allow and prevent bypass air flow therethrough.

9. The compound rotorcraft as recited in claim 1 wherein the fan module further comprises a plurality of inlet guide vanes actuatable between open and closed positions to allow and prevent bypass air flow therethrough.

10. The compound rotorcraft as recited in claim 1 wherein the fan module is positioned proximate an aft portion of the core turboshaft engine.

11. The compound rotorcraft as recited in claim 1 wherein the thrust nozzle is selected from the group consisting of a fixed nozzle, a moveable nozzle, a thrust vectoring nozzle and an axisymmetric convergent/divergent nozzle.

12. The compound rotorcraft as recited in claim 1 wherein, in the turboshaft and turbofan configuration, the hybrid propulsion engine further comprises a bypass ratio of between about 1 to 1 and about 4 to 1.

13. The compound rotorcraft as recited in claim 1 wherein, in the turboshaft and turbofan configuration, the hybrid propulsion engine further comprises a bypass ratio of between about 2 to 1 and about 3 to 1.

14. The compound rotorcraft as recited in claim 1 wherein, in the turboshaft and turbofan configuration, the hybrid propulsion engine further comprises a bypass ratio of about 2.5 to 1.

15. The compound rotorcraft as recited in claim 1 wherein, in the turboshaft configuration, power from the core turboshaft engine is directed to the main rotor.

16. The compound rotorcraft as recited in claim 1 wherein, in the turboshaft and turbofan configuration, power from the core turboshaft engine is shared by the main rotor and the fan module.

Assignees

• Textron Innovations Inc

Inventors

• Ross Brent Chadwick
• Ivans Steven Ray
• Ewing Alan Carl

Classifications

• B64D27/02

  Aircraft characterised by the type or position of power plants · CPC title

• F02C7/042

  having variable geometry · CPC title

• F02K1/002

  with means to modify the direction of thrust vector (F02K1/54 takes precedence; thrust vectoring of rockets F02K9/80) · CPC title

• F02K3/062

  with aft fan · CPC title

• F02K3/075

  controlling flow ratio between flows · CPC title