ISOLATED TURBINE ENGINE COOLING

• US 20200131933A1
• Filed: 10/29/2018
• Published: 04/30/2020
• Est. Priority Date: 10/29/2018
• Status: Active Grant

• Current Assignee: Rolls-Royce North American Technologies, Inc. (Rolls-Royce Holdings Plc)
• Original Assignee: Rolls-Royce North American Technologies, Inc. (Rolls-Royce Holdings Plc)
• Inventors: Metzger, Nicholas Dwain
• Granted Patent: US 10,927,708 B2
• Time in Patent Office: Days
• CPC Class Codes: B64C 30/00 Supersonic type aircraft
• CPC Class Codes: B64D 2033/024 comprising cooling means
• CPC Class Codes: B64D 2033/026 for supersonic or hypersonic aircraft
• CPC Class Codes: B64D 27/026 comprising different types of power plants, e.g. combination of a piston engine and a gas-turbine Hybrid electric aircraft B64D27/33
• CPC Class Codes: B64D 27/10 of gas-turbine type jet aircraft B64D27/16
• CPC Class Codes: B64D 27/14 within, or attached to, fuselages
• CPC Class Codes: B64D 27/16 of jet type
• CPC Class Codes: B64D 27/20 within, or attached to, fuselages
• CPC Class Codes: B64D 33/04 of exhaust outlets or jet pipes exhaust outlets for combustion engines in general F01N; jet pipes or nozzles for jet-propulsion plants per se F02K; plants characterised by the form or arrangement of the jet pipe or nozzle F02K ; attitude, flight direction, or altitude control by jet reaction B64C
• CPC Class Codes: F01D 21/00 Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for
• CPC Class Codes: F01D 25/145 Thermally insulated casings
• CPC Class Codes: F02C 7/04 Air intakes for gas-turbine plants or jet-propulsion plants
• CPC Class Codes: F02C 7/12 Cooling of plants of component parts, see the relevant subclasses, e.g. F01D; cooling of engines in general F01P
• CPC Class Codes: F02K 7/16 Composite ram-jet/turbo-jet engines
• CPC Class Codes: F05D 2240/14 Casings or housings protecting or supporting assemblies within
• CPC Class Codes: F05D 2240/15 Heat shield
• CPC Class Codes: F05D 2260/232 characterized by the cooling medium
• CPC Class Codes: F05D 2270/112 by limiting temperatures
• CPC Class Codes: F05D 2270/303 Temperature
• CPC Class Codes: Y02T 50/60 Efficient propulsion technologies, e.g. for aircraft

Abstract

A hybrid propulsion system and methods for cooling the same are provided. The system may comprise a gas turbine and a secondary engine. The gas turbine engine may have a core passage and an engine compartment. The secondary engine may be a supersonic and/or hypersonic engine. The system may comprise a thermal barrier, an inlet and an exhaust. The thermal barrier may longitudinally envelope the gas turbine engine. The thermal barrier may comprise an inner envelope, an outer envelope, an upstream opening, and a downstream opening. The inlet may be in fluid communication with the ambient environment and the gas turbine engine via the upstream opening. The exhaust may be in fluid communication with the ambient environment and the gas turbine engine via the downstream opening. The engine compartment may be located between a boundary of the core passage and the inner envelope.

Claims

1. A hybrid propulsion system comprising:

a gas turbine engine, the gas turbine having a core passage and an engine compartment, the core passage including a compressor, combustion chamber and a turbine;

a hypersonic secondary engine;

a thermal barrier longitudinally enveloping the gas turbine engine, the thermal barrier including;

an inner envelope and an outer envelope;

a vacuum between the inner envelope and the outer envelope; and

an upstream opening and a downstream opening;

an inlet in fluid communication with the ambient environment and the gas turbine engine via at least the upstream opening; and

an exhaust in fluid communication with the ambient environment and the gas turbine engine via at least the downstream opening,

wherein the engine compartment is located between a boundary of the core passage and the inner envelope and extending at least between the upstream opening and the downstream opening.

2. The system of claim 1, further comprising an inlet flow deflector limiting ambient air into the engine compartment.

3. The system of claim 1, further comprising an exhaust flow deflector restricting flow in the engine compartment.

4. The system of claim 1, further comprising a plurality of nozzles within the engine compartment proximate to the upstream opening.

5. The system of claim 4, further comprising a cooling fluid source connected to the plurality of nozzles.

6. The system of claim 1, further comprising a plurality of mixing channels extending between the thermal barrier and the core passage at the upstream opening.

7. The system of claim 1, further comprising a plurality of nozzles within the core passage, the plurality of nozzles connected to a source to provide a cooling fluid.

8. The system of claim 7, wherein the cooling fluid is selected from the group consisting of bleed air, water, fuel and refrigerant.

9. The system of claim 1, wherein the hypersonic secondary engine comprises a ramjet engine or scramjet engine.

10. The system of claim 1, wherein the inner envelop and outer envelope are concentric cylinders.

11. A hybrid propulsion system comprising:

a gas turbine engine including a core passage having a compressor, combustion chamber and a turbine;

a hypersonic secondary engine;

a thermal barrier longitudinally enveloping the gas turbine engine, the thermal barrier including;

an inner envelope and an outer envelope;

a passage between the inner envelope and the outer envelope;

an upstream opening and a downstream opening; and

a cooling fluid inlet at proximate downstream opening and a cooling fluid exit nozzle proximate to the upstream opening;

the cooling fluid inlet in fluid communication with the cooling fluid exit nozzle via the passage;

an inlet in fluid communication with the ambient environment and the gas turbine engine via at least the upstream opening; and

an exhaust in fluid communication with the ambient environment and the gas turbine engine via at least the downstream opening.

12. The system of claim 11, wherein the gas turbine engine further comprises an engine compartment located between a boundary of the core passage and the inner envelope and extending at least between the upstream opening and the downstream opening.

13. The system of claim 12, further comprising a plurality of nozzles within the engine compartment.

14. The system of claim 13, wherein the plurality of nozzles are connected to the passage.

15. The system of claim 12, further including a source of cooling fluid, wherein the cooling fluid is selected from the group consisting of bleed air, water, fuel or refrigerant.

16. The system of claim 12, further comprising a plurality of heat exchanger supplied with cooling bleed air at the upstream opening and prior to a plurality mixing channels extending between the thermal barrier and the core passage at the upstream opening.

17. The system of claim 11, further comprising an expansion device, a condenser with a heat exchanger, a fuel line and a pump, and a closed flow path defined by the passage, through the expansion device, the condenser, the pump and back to the passage;

wherein the fuel line is in fluid communication with the heat exchanger.

18. A method of cocooning a gas turbine engine within a hypersonic aircraft comprising:

isolating the gas turbine engine from an airframe of the aircraft with a double walled barrier, the double walled barrier defining a passage between inner and outer walls;

providing thrust to the aircraft via the gas turbine engine, wherein the gas turbine engine receives ambient air though an inlet to the gas turbine, compresses the air in a compressor, heats the air in a combustion chamber and expands the air in a turbine and exhausts the air from an exhaust to the environment;

switching from the gas turbine engine to a secondary propulsion device;

restricting air flow through the gas turbine engine;

supplying a cooling fluid into the passage; and

spraying cooling fluid from the passage into the gas turbine.

19. The method of claim 18, wherein the step of restricting air flow comprises moving a deflector in the inlet to a closed position and moving a second deflector in the exhaust to a closed position.

20. The method of claim 18, regulating the supply of cooling fluid to into the fluid passage as a function of the core temperature of the gas turbine engine.