Systems and methods for hybrid electric gas turbine engines
AssigneeRaytheon Technologies Corporation
InventorsDaniel Kupratis, Gary Roberge
Life of this patent
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Abstract
A hybrid electric gas turbine propulsion system may comprise: a first propulsion system, a second propulsion system, and a third propulsion system. The first propulsion system may comprise a first fan, a first turbine, a first compressor, and a first electric motor, the first fan operably coupled to the first turbine and the first compressor by a first shaft, the first shaft coupled to the first electric motor, the first shaft configured to be disposed radially inward of a fuselage of an aircraft. The second propulsion system and the third propulsion system may be in accordance with the first propulsion system. The hybrid electric gas turbine propulsion system may be symmetric about a vertical plane extending through a neutral aerodynamic axis.
Claims20
A hybrid electric gas turbine propulsion system, comprising:
- a first propulsion system including a first fan, a first turbine, a first compressor, and a first electric motor, the first fan operably coupled to the first turbine and the first compressor by a first shaft, the first shaft coupled to the first electric motor, and the first shaft configured to be disposed radially inward of a fuselage of an aircraft;
- a second propulsion system including a second fan, a second turbine, a second compressor, and a second electric motor, the second fan operably coupled to the second turbine and the second compressor by a second shaft, the second shaft coupled to the second electric motor, and the second shaft configured to be disposed radially inward of the fuselage of the aircraft;
- a first splitter at least partially defining a first inlet of the first fan, the first inlet in fluid communication with the first fan;
- and
- a second splitter at least partially defining a second inlet of the first compressor, the second splitter disposed aft of the first splitter, the second splitter configured to split a turbine discharge flow.
A boundary layer impulse propulsion system, comprising:
- a fan section disposed radially inward of a fuselage of an aircraft, the fan section including a plurality of fans disposed circumferentially about a neutral aerodynamic axis of the fuselage, and the fan section configured to ingest a first boundary layer airflow from the fuselage, the plurality of fans including at least three fans, each fan in the plurality of fans coupled to a corresponding shaft of a plurality of shafts;
- a turbine section disposed aft of the fan section, the turbine section including a plurality of turbines, the plurality of turbines including at least three turbines;
- and
- a compressor section disposed aft of the turbine section, the compressor section including a plurality of compressors, the compressor section configured to ingest a second boundary layer flow from the fuselage, and each compressor in the compressor section being in fluid communication with a respective turbine in the turbine section.
An aircraft, comprising:
- a fuselage extending from a forward end to an aft end, the fuselage defining a neutral aerodynamic axis;
- and
- a plurality of propulsion systems disposed proximate the aft end of the fuselage, wherein:
- at least a portion of each propulsion system in the plurality of propulsion systems is disposed inwardly of the fuselage, wherein:
- each propulsion system in the plurality of propulsion systems is configured to ingest a boundary layer flow from the fuselage,
- a first fan of a first propulsion system in the plurality of propulsion systems mirrors a second fan of a second propulsion system in the plurality of propulsion systems about a plane extending through the neutral aerodynamic axis, and a longitudinal axis of a propulsion system in the plurality of propulsion systems, and
- the first fan, the second fan, and a third fan of a third propulsion system in the plurality of propulsion systems are disposed circumferentially about the neutral aerodynamic axis of the fuselage.
Dependencies are inferred from the claim text.