Abstract: A gas turbine engine is provided. The gas turbine engine includes a turbomachine defining an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct; a primary fan driven by the turbomachine; and a secondary fan located downstream of the primary fan within the inlet duct. The gas turbine engine defines a thrust to power airflow ratio between 3.5 and 100 and a core bypass ratio between 0.1 and 10, wherein the thrust to power airflow ratio is a ratio of an airflow through a bypass passage over the turbomachine plus an airflow through the fan duct to an airflow through the core duct, and wherein the core bypass ratio is a ratio of the airflow through the fan duct to the airflow through the core duct.

Abstract: A gas turbine engine includes a turbomachine defining an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct, a primary fan driven by the turbomachine, a secondary fan located downstream of the primary fan within the inlet duct, a heat exchanger disposed in the fan duct, and a booster upstream of the heat exchanger, the booster including a booster cowl extending into the fan duct, the booster cowl separating an upstream portion of the fan duct into an upper fan duct having an upper fan duct inlet and a lower fan duct having a lower fan duct inlet, the upper fan duct inlet and lower fan duct inlet collectively forming the fan duct inlet.

Abstract: A method is provided of generating electric power with an electric machine. The method includes rotating a rotor of the electric machine relative to a stator of the electric machine with a shaft of a gas turbine engine during an operating condition of the gas turbine engine, the gas turbine engine being a three-stream gas turbine engine defining an axial direction, the three-stream gas turbine engine comprising: the shaft, a primary fan operatively coupled with the shaft, a mid-fan positioned downstream of the primary fan and operatively coupled with the shaft, a low pressure turbine operatively coupled with the shaft, wherein rotating the rotor of the electric machine relative to the stator of the electric machine comprises generating an electric machine power during the operating condition and generating a low pressure turbine power during the operating condition.

Abstract: A gas turbine engine is provided, comprising: a turbomachine having a compressor section, a combustion section, and a turbine section in serial flow order; and a fan section having a fan drivingly coupled to the turbomachine and an airflow surface rotatable with the fan and exposed to a fan airflow provided to and through the fan during operation of the gas turbine engine, the airflow surface defining a plurality of boundary layer openings configured to ingest a boundary layer of the fan airflow over the airflow surface during operation of the gas turbine engine.

Abstract: A propulsion system, the propulsion system comprising a rotating element, a stationary element, and an inlet between the rotating element and the stationary element, wherein the inlet passes radially inward of the stationary element; wherein the inlet passes radially inward of the stationary element; wherein the inlet leads to an inlet duct containing a ducted fan having an axis of rotation and a plurality of blades; and wherein the inlet duct divides into a first duct and a second duct, separate from the first duct. A method of operating a propulsion system, comprising the steps of: operating a first rotating fan assembly to produce a first stream of air; directing a portion of the first stream of air into a second ducted rotating fan assembly; operating the second ducted rotating fan assembly to produce a second stream of air; dividing the second stream of air into a core stream and a fan stream; and directing the core stream into a gas turbine engine core.

Abstract: A method is provided for operating a hybrid-electric propulsion system having a first engine, a second engine, a first electric machine coupled to the first engine, and a second electric machine coupled to one of the first engine or the second engine. The method includes: receiving data indicative of a first engine operating parameter, a second engine operating parameter, or both; determining a first engine operating parameter margin, a second parameter operating margin, or both; determining a load share for the first engine, the second engine, or both, or between the first engine and the second engine based on the first engine operating parameter margin, the second engine operating parameter margin, or both; and transferring a first amount of power to or from the first electric machine and a second amount of power to or from the second electric machine in response to the determined load share.

Abstract: A gas turbine engine is provided. The gas turbine engine includes a turbomachine defining an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct; a primary fan driven by the turbomachine; and a secondary fan located downstream of the primary fan within the inlet duct. The gas turbine engine defines a thrust to power airflow ratio between 3.5 and 100 and a core bypass ratio between 0.1 and 10, wherein the thrust to power airflow ratio is a ratio of an airflow through a bypass passage over the turbomachine plus an airflow through the fan duct to an airflow through the core duct, and wherein the core bypass ratio is a ratio of the airflow through the fan duct to the airflow through the core duct.

Abstract: A gas turbine engine is provided. The gas turbine engine includes a turbomachine defining an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct; a primary fan driven by the turbomachine; and a secondary fan located downstream of the primary fan within the inlet duct. The gas turbine engine defines a thrust to power airflow ratio between 3.5 and 100 and a core bypass ratio between 0.1 and 10, wherein the thrust to power airflow ratio is a ratio of an airflow through a bypass passage over the turbomachine plus an airflow through the fan duct to an airflow through the core duct, and wherein the core bypass ratio is a ratio of the airflow through the fan duct to the airflow through the core duct.

Abstract: A gas turbine is provided, the gas turbine engine including a turbomachine having an inlet splitter defining in part an inlet to a working gas flowpath and a fan duct splitter defining in part an inlet to a fan duct flowpath. The gas turbine engine also includes a primary fan driven by the turbomachine defining a primary fan tip radius R1, a primary fan hub radius R2, and a primary fan specific thrust rating TP; and a secondary fan downstream of the primary fan and driven by the turbomachine, the secondary fan defining a secondary fan tip radius R3, a secondary fan hub radius R4, and a secondary fan specific thrust rating TS; wherein the gas turbine engine defines an Effective Bypass Area, and wherein a ratio of R1 to R3 equals R 1 R 3 = ( EFP ) ? ( 1 - RqR ( Sec . - Fan ) 2 ) ( 1 - RqR Prim . - Fan 2 ) ? ( T P T S ) ? ( EBA ) .

Abstract: A propulsion system, the propulsion system comprising a rotating element, a stationary element, and an inlet between the rotating element and the stationary element, wherein the inlet passes radially inward of the stationary element; wherein the inlet passes radially inward of the stationary element; wherein the inlet leads to an inlet duct containing a ducted fan having an axis of rotation and a plurality of blades; and wherein the inlet duct divides into a first duct and a second duct, separate from the first duct. A method of operating a propulsion system, comprising the steps of: operating a first rotating fan assembly to produce a first stream of air; directing a portion of the first stream of air into a second ducted rotating fan assembly; operating the second ducted rotating fan assembly to produce a second stream of air; dividing the second stream of air into a core stream and a fan stream; and directing the core stream into a gas turbine engine core.

Abstract: A gas turbine engine is provided. The gas turbine engine includes a turbomachine defining an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct; a primary fan driven by the turbomachine; and a secondary fan located downstream of the primary fan within the inlet duct. The gas turbine engine defines a thrust to power airflow ratio between 3.5 and 100 and a core bypass ratio between 0.1 and 10, wherein the thrust to power airflow ratio is a ratio of an airflow through a bypass passage over the turbomachine plus an airflow through the fan duct to an airflow through the core duct, and wherein the core bypass ratio is a ratio of the airflow through the fan duct to the airflow through the core duct.

Abstract: A method is provided for operating a hybrid-electric propulsion system having a first engine, a second engine, a first electric machine coupled to the first engine, and a second electric machine coupled to one of the first engine or the second engine. The method includes: receiving data indicative of a first engine operating parameter, a second engine operating parameter, or both; determining a first engine operating parameter margin, a second parameter operating margin, or both; determining a load share for the first engine, the second engine, or both, or between the first engine and the second engine based on the first engine operating parameter margin, the second engine operating parameter margin, or both; and transferring a first amount of power to or from the first electric machine and a second amount of power to or from the second electric machine in response to the determined load share.

Abstract: A propulsion system is provided including an unducted rotating fan defining a fan axis; and a turbomachine disposed downstream from the unducted rotating fan, wherein the turbomachine defines a working gas flowpath flowing therethrough; wherein the propulsion system defines a third stream flowpath and an inlet passage having an inlet that is offset from the fan axis, wherein the inlet passage is configured to provide an inlet airflow to the working gas flowpath, and wherein the third stream flowpath bypasses at least a portion of the turbomachine.

Abstract: An electric propulsor is provided. The electric propulsor includes a core cowl and an outer cowl. A first air flowpath is defined radially outward of the outer cowl and a second air flowpath is defined between the core cowl and the outer cowl. The electric propulsor also includes one or more electric machines, a fan rotatably drivable by at least one electric machine, and a booster having a plurality of airfoils disposed at least in part in the second air flowpath, the booster being rotatably drivable by at least one electric machine for compressing air flowing along the second air flowpath. The electric propulsor further includes a heat exchanger disposed within the second air flowpath downstream of the booster, the heat exchanger being in thermal communication with at least one of the one or more electric machines and/or a gearbox mechanically coupled with the fan.

Abstract: A gas turbine engine is provided. The gas turbine engine includes a turbomachine defining an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct; a primary fan driven by the turbomachine; and a secondary fan located downstream of the primary fan within the inlet duct. The gas turbine engine defines a thrust to power airflow ratio between 3.5 and 100 and a core bypass ratio between 0.1 and 10, wherein the thrust to power airflow ratio is a ratio of an airflow through a bypass passage over the turbomachine plus an airflow through the fan duct to an airflow through the core duct, and wherein the core bypass ratio is a ratio of the airflow through the fan duct to the airflow through the core duct.

Abstract: A gas turbine engine defining a centerline and a circumferential direction, the gas turbine engine including: a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the turbomachine defining a working gas flowpath and a fan duct flowpath; a primary fan driven by the turbomachine defining a primary fan tip radius R1 and a primary fan hub radius R2; a secondary fan located downstream of the primary fan and driven by the turbomachine, at least a portion of an airflow from the primary fan configured to bypass the secondary fan, the secondary fan defining a secondary fan tip radius R3 and a secondary fan hub radius R4, wherein the secondary fan is configured to provide a fan duct airflow through the fan duct flowpath during operation to generate a fan duct thrust, wherein the fan duct thrust is equal to % Fn3S of a total engine thrust during operation of the gas turbine engine at a rated speed during standard day operating conditions; wherein a ra

Abstract: A gas turbine is provided, the gas turbine engine including a turbomachine having an inlet splitter defining in part an inlet to a working gas flowpath and a fan duct splitter defining in part an inlet to a fan duct flowpath. The gas turbine engine also includes a primary fan driven by the turbomachine defining a primary fan tip radius R1, a primary fan hub radius R2, and a primary fan specific thrust rating TP; and a secondary fan downstream of the primary fan and driven by the turbomachine, the secondary fan defining a secondary fan tip radius R3, a secondary fan hub radius R4, and a secondary fan specific thrust rating TS; wherein the gas turbine engine defines an Effective Bypass Area, and wherein a ratio of R1 to R3 equals R 1 R 3 = ( E ? F ? P ) ? ( 1 - RqR Sec . - Fan 2 ) ( 1 - RqR Prim . - Fan 2 ) ? ( T P T S ) ? ( E ? B ? A ) .

Abstract: A method for operating a hybrid-electric propulsion system of an aircraft, the hybrid-electric propulsion system, the method comprising: sensing data indicative of at least one of an aerodynamic instability, a pressure, or a temperature within the HP compressor and the LP compressor of the gas turbine engine; identifying a aerodynamically unstable compressor by determining that conditions within one of the HP compressor or the LP compressor are within a threshold of a stall condition based at least in part on the sensed data within the HP compressor and the LP compressor of the gas turbine engine, the one of the HP compressor or the LP compressor that is determined to be within the threshold of a stall condition being the aerodynamically unstable compressor; and transferring power, via the one or more electric machines, to the aerodynamically unstable compressor in order to clear the stall condition.

Abstract: A gas turbine is provided, the gas turbine engine including a turbomachine having an inlet splitter defining in part an inlet to a working gas flowpath and a fan duct splitter defining in part an inlet to a fan duct flowpath. The gas turbine engine also includes a primary fan driven by the turbomachine defining a primary fan tip radius R1, a primary fan hub radius R2, and a primary fan specific thrust rating TP; and a secondary fan downstream of the primary fan and driven by the turbomachine, the secondary fan defining a secondary fan tip radius R3, a secondary fan hub radius R4, and a secondary fan specific thrust rating TS; wherein the gas turbine engine defines an Effective Bypass Area, and wherein a ratio of R1 to R3 equals R 1 R 3 = ( EFP ) ? ( 1 - RqR ( Sec . - Fan ) 2 ) ( 1 - RqR Prim . - Fan 2 ) ? ( T P T S ) ? ( EBA ) .

Abstract: A gas turbine is provided, the gas turbine engine including a turbomachine having an inlet splitter defining in part an inlet to a working gas flowpath and a fan duct splitter defining in part an inlet to a fan duct flowpath. The gas turbine engine also includes a primary fan driven by the turbomachine defining a primary fan tip radius R1, a primary fan hub radius R2, and a primary fan specific thrust rating TP; and a secondary fan downstream of the primary fan and driven by the turbomachine, the secondary fan defining a secondary fan tip radius R3, a secondary fan hub radius R4, and a secondary fan specific thrust rating TS; wherein the gas turbine engine defines an Effective Bypass Area, and wherein a ratio of R1 to R3 equals R 1 R 3 = ( E ? F ? P ) ? ( 1 - RqR Sec . - Fan 2 ) ( 1 - RqR Prim . - Fan 2 ) ? ( T P T S ) ? ( E ? B ? A ) .