Dual mode turbofan engine

Patent metadata

Field	Value
Publication number	US-10711791-B1
Application number	US-201816103291-A
Country	US
Kind code	B1
Filing date	Aug 14, 2018
Priority date	Apr 1, 2014
Publication date	Jul 14, 2020
Grant date	Jul 14, 2020

Abstract

A dual mode turbofan engine includes a jet engine portion having a compressor, a turbine disposed aft of the compressor, and a shaft coupled to the compressor and the turbine. The jet engine portion is configured to produce an exhaust. The system further includes an auxiliary turbine having a plurality of auxiliary turbine blades. The auxiliary turbine is disposed aft of the turbine and decoupled from the shaft. The system also includes a diverter disposed between the turbine and the auxiliary turbine. The diverter is configured to selectively direct the exhaust to an inner flow path bypassing the plurality of auxiliary turbine blades or to an outer flow path engaging the plurality of turbine blades. A plurality of propeller blades is hingedly connected to the auxiliary turbine.

First claim

What is claimed is:

1. A dual mode turbofan engine system, the engine system comprising:
a jet engine portion having a compressor, a turbine disposed aft of the compressor, and a shaft coupled to the compressor and the turbine, wherein the jet engine portion is configured to produce an exhaust;
an auxiliary turbine having a plurality of auxiliary turbine blades, wherein the auxiliary turbine is disposed aft of the turbine and decoupled from the shaft;
a diverter disposed between the turbine and the auxiliary turbine, wherein the diverter is configured to selectively direct the exhaust to an inner flow path bypassing the plurality of auxiliary turbine blades or to an outer flow path engaging the plurality of turbine blades; and
a plurality of propeller blades hingedly connected to the auxiliary turbine;
wherein the diverter is an adjustable frustoconical structure, and the diverter includes a plurality of overlapping plates.

2. The system of claim 1 , wherein each of the plurality of propeller blades are spring biased to a swept position, wherein said spring bias is overcome when the auxiliary turbine is actuated.

3. The system of claim 1 , wherein each of the plurality of propeller blades are selectably lockable in either a retracted or an extended position.

4. The system of claim 1 , wherein the diverter includes a fluidic structure configured to direct the exhaust to the inner flow path or to the outer flow path.

5. The system of claim 4 , wherein the fluidic structure is a ring or nozzle ring having an exhaust annulus oriented radially inward or radially outward from the ring.

6. A dual mode turbofan engine system, the engine system comprising:
a jet engine portion having a compressor, a turbine disposed aft of the compressor, and a shaft coupled to the compressor and the turbine, wherein the jet engine portion is configured to produce an exhaust;
an auxiliary turbine having a plurality of auxiliary turbine blades, wherein the auxiliary turbine is disposed aft of the turbine and decoupled from the shaft;
a diverter disposed between the turbine and the auxiliary turbine, wherein the diverter is configured to selectively direct the exhaust to an inner flow path bypassing the plurality of auxiliary turbine blades or to an outer flow path engaging the plurality of turbine blades;
a generator or an alternator operatively coupled to the auxiliary turbine; and
a remotely disposed motor having cooperating propeller blades, wherein the remotely disposed motor is electrical communication with the alternator or the generator;
wherein the diverter is an adjustable frustoconical structure including a plurality of overlapping plates.

7. The system of claim 6 , wherein the diverter includes a fluidic structure configured to direct the exhaust to the inner flow path or to the outer flow path.

8. The system of claim 7 , wherein the fluidic structure is a ring or nozzle ring having an exhaust annulus oriented radially inward or radially outward from the ring.

9. The system of claim 7 , wherein the compressor includes a stator having an aft end, and wherein the fluidic structure is a plurality of exit guide vanes disposed proximate the aft end of the stator.

Assignees

• Us Gov Air Force
• United States As Represented By The Secretary Of The Air Force

Inventors

• Rolling August J
• Wood Jeffrey K

Classifications

• B64U50/15

  using combustion exhausts other than turbojets or turbofans, e.g. using rockets, ramjets, scramjets or pulse-reactors · CPC title

• B64U30/12

  Variable or detachable wings, e.g. wings with adjustable sweep · CPC title

• B64U50/12

  using turbine engines, e.g. turbojets or turbofans · CPC title

• B64U50/19

  using electrically powered motors · CPC title

• B64U50/13

  using external fans or propellers · CPC title