Patent Description:
An aircraft generally includes a propulsion system that provides thrust. The propulsion system can include at least two aircraft engines. Each engine is typically mounted to a respective one of the wings of the aircraft or at other practicable locations. While a gas turbine engine is lighter and can produce more thrust than an internal combustion engine, the internal combustion engine may have better fuel burn characteristics. Accordingly, a propulsion system capable of operating utilizing the particular benefits of each type of engine would be useful.

<CIT> discloses an aircraft propulsion system which includes a first gas turbine engine including a first input shaft driving a first gear system, a first fan driven by the first gear system, a first generator supported on the first input shaft and a fan drive electric motor providing a drive input to the first fan, a second gas turbine engine including a second input shaft driving a second gear system, a second fan driven by the second gear system, a second generator supported on the second input shaft and a second fan drive electric motor providing a drive input to the second fan and a controller controlling power output from each of the first and second generators and directing the power output between each of the first and second fan drive electric motors.

<CIT> discloses a vertical take-off and landing aircraft including a wing structure including a wing, a rotor operatively supported by the wing, and a hybrid power system configured to drive the rotor, the hybrid power system including a first power system and a second power system, wherein a first energy source for the first power system is different than a second energy source for the second power system.

In some embodiments of the present disclosure, a hybrid propulsion system as claimed in claim <NUM> is provided.

In some embodiments of the present disclosure, a method as claimed in claim <NUM> is provided.

These and other features, aspects, and advantages of the present invention will become better understood with reference to the following description and appended claims.

Accordingly, a value modified by a term or terms, such as "about", "approximately", "generally", and "substantially", are not to be limited to the precise value specified.

For example, if a composition or assembly is described as containing components A, B, and/or C, the composition or assembly can contain A alone; B alone; C alone; A and B in combination; A and C in combination; B and C in combination; or A, B, and C in combination.

A hybrid propulsion system for an aircraft is provided herein that includes a propulsor assembly having at least one propulsor and a power generation assembly. The power generation assembly includes a first power assembly, a second power assembly, a first electric machine, and a second electric machine. The first power assembly is drivingly coupled to the first electric machine to produce a first amount of electric power. The second power assembly is drivingly coupled to the second electric machine to produce a second amount of electric power, wherein the second power assembly is configured to generate electric power more efficiently than the first power assembly. The propulsion system further comprises a power bus coupled to the first electric machine and the second electric machine.

A controller is operably coupled to the first power assembly, the first electric machine, or both and to the second power assembly, the second power assembly, or both, the controller and the power bus configured to combine at least a portion of the first and second amount of power for electric transfer to the propulsor assembly. The first power assembly includes a turbomachine and the second power assembly includes an internal combustion engine. The controller can also be configured to provide electrical power from the first electric machine to the propulsor assembly in a first operating condition and from the second electric machine to the propulsor assembly in a second operating condition. In some instances, the first operating condition may be indicated by a command to accelerate or climb the aircraft and the second operating condition may be indicated by a command to operate the aircraft in a level flight condition. Additionally, or alternatively, in various embodiments, the controller can be configured to receive a desired thrust output, and if a desired thrust output is within a second power assembly operating range, activating the second power assembly, and if the desired thrust output is greater than the second power assembly operating range, activating the first power assembly.

In some instances, the first power assembly may be deactivated when the aircraft is in the second operating condition and the second power assembly may be deactivated when the aircraft is in the first operating condition. Additionally, or alternatively, one of the first and second propulsor assemblies may be utilized to provide a desired amount of thrust for the aircraft while the other of the first and second power assembly may simultaneously be used for generating electric power for one or more power loads of the aircraft.

In various embodiments, the first electric machine and the second electric machine are both configured to generate electrical power that is stored within the energy storage unit. In some cases, the second electric machine may be configured to generate less electrical power than the first electric machine. For instance, the second electric machine can generate less than half of the electrical power of the first electric machine.

In some embodiments, an energy storage unit can be operably coupled with each of the first power assembly, the second power assembly, and the propulsor assembly. Each of the first power assembly, the second power assembly, and the propulsor assembly can be configured to utilize electrical power stored in the energy storage unit.

By operating in accordance with one or more these aspects, the hybrid propulsion system provided herein may provide a sufficient amount of thrust output to the aircraft at each operating condition of the aircraft (e.g., takeoff, cruise, loiter, etc.) while operating in an efficient manner. For instance, the hybrid propulsion system provided herein may use a first power assembly that incorporates a lightweight, high specific power engine (e.g., a constant combustion engine) for takeoff and/or dash conditions and a second power assembly that incorporates a high efficiency, low specific power engine (e.g., an intermittent combustion engine) for long-duration cruise/loiter conditions. In addition, the hybrid propulsion system may further incorporate a propulsor assembly that uses a partial or full electrical drive-train that can enable the combination of power from disparate plants in order to drive one or more propulsors. The one or more propulsors may be positioned in any practicable location about the aircraft. In various embodiments, through one or more computing systems, various configurations of thrust outputs may occur when each of the one or more propulsors is switched on or off, based on the system power demand, and the propulsor assembly can compensate for the hybrid propulsion system as one power assembly is brought up to speed during a power transition. In some embodiments, the first power assembly may be configured as a gas turbine that is lighter than the second power assembly based on a weight/pounds of thrust capable of being produced ratio and can produce more electric power and/or thrust than the second power assembly. Conversely, when the second power assembly is configured as an internal combustion engine, it may be heavier than the first power assembly based on a weight/pounds of thrust capable of being produced ratio, but may have better fuel burn characteristics than the first power assembly. By creating a hybrid propulsion system that utilizes the first power assembly for the high power take-off or high-speed conditions but relies (possibly solely) upon the second power assembly for cruise or loiter conditions may increase the aircraft capability while improving the overall mission duration. In some instances, the hybrid propulsion system provided herein may lead to double digit range improvements for a multi-gas turbine architecture and range doubling for a diesel-gas turbine architecture when compared to a single gas turbine architecture.

A system architecture that combines multiple engine types and uses each engine for the part of the mission where it is most efficient results in increased performance and reduced fuel burn throughout the mission. The amount of benefit from the combined architecture is based on the duration aloft, making the disclosed architecture beneficial for prolonged endurance applications. Any extended capability and reduced fuel burn is both a commercial and military advantage over currently available propulsion systems.

Referring now to the drawings, wherein identical numerals indicate the same elements throughout the figures, <FIG> generally provides a schematic view of an aircraft <NUM> that may incorporate various features of the present disclosure. As shown in <FIG>, the aircraft <NUM> defines a longitudinal centerline <NUM> that extends therethrough, a lateral direction L, a forward end portion <NUM>, and an aft end portion <NUM>. Moreover, the aircraft <NUM> includes a fuselage <NUM>, extending longitudinally from the forward end portion <NUM> of the aircraft <NUM> to the aft end portion <NUM> of the aircraft <NUM>, and an empennage <NUM> at the aft end portion of the aircraft <NUM>.

The aircraft <NUM> can also include a wing assembly including a first, port side wing <NUM> and a second, starboard side wing <NUM>. The first and second wings <NUM>, <NUM> each extend laterally outward with respect to the longitudinal centerline <NUM>. The first wing <NUM> and a portion of the fuselage <NUM> together define a first side <NUM> of the aircraft <NUM>, and the second wing <NUM> and another portion of the fuselage <NUM> together define a second side <NUM> of the aircraft <NUM>. For the embodiment depicted, the first side <NUM> of the aircraft <NUM> is configured as the port side of the aircraft <NUM>, and the second side <NUM> of the aircraft <NUM> is configured as the starboard side of the aircraft <NUM>.

Each of the wings <NUM>, <NUM> for the embodiment depicted includes one or more flaps <NUM>, which may be in the form of leading-edge flaps and one or more trailing-edge flaps. The aircraft <NUM> further includes, or rather, the empennage <NUM> of the aircraft <NUM> includes, a vertical stabilizer <NUM> which may have a rudder flap for yaw control, and a pair of horizontal stabilizers <NUM>, each having an elevator flap <NUM> for pitch control. The fuselage <NUM> additionally includes an outer surface or skin <NUM>. It will be appreciated, however, that in other embodiments of the present disclosure, the aircraft <NUM> may additionally or alternatively include any other suitable configuration. For example, in other embodiments, the aircraft <NUM> may include any other configuration of stabilizer. Moreover, it will be appreciated that in some embodiments, the aircraft <NUM> may be configured as a vertical takeoff and landing (VTOL) aircraft, a helicopter, or any other type of aerial vehicle without departing from the scope of the present disclosure.

In some embodiments, the aircraft <NUM> may be an unmanned aerial vehicle capable of flight without a human pilot aboard. For example, the aircraft <NUM> may be piloted by, e.g., remote control by a human operator, or alternatively, may be fully or intermittently autonomous and controlled by onboard and/or offboard computers.

Referring to <FIG>, the aircraft <NUM> of <FIG> can include a hybrid propulsion system <NUM> having a power generation assembly, which can include a first power assembly <NUM> and/or a second power assembly <NUM>, and one or more propulsor assemblies <NUM>. As used herein, "hybrid" can indicate any propulsion system that includes more than one type of power source. For instance, in various embodiments, the power generation assembly can include turbine engines (e.g., continuous combustion or continuous rotation engines), internal combustion engines (e.g., intermittent combustion or reciprocating engines), electric machines and/or any other type of machine that can generate electrical power and/or propulsive force for the aircraft <NUM>. For example, <FIG> provides a schematic, cross-sectional view of a first power assembly <NUM>, <FIG> provides a schematic, cross-sectional view of a second power assembly <NUM>, and <FIG> provides a schematic, cross-sectional view of a propulsor assembly <NUM> that can be operably coupled with the first and/or the second power assembly <NUM>, <NUM>. In various embodiments, each of the first and/or second power assemblies <NUM>, <NUM> may each be configured in an underwing-mounted configuration and produce electrical power and/or thrust in varied manners. Additionally, and/or alternatively, the first and/or the second power assemblies <NUM>, <NUM> may be operably coupled with the fuselage of the aircraft <NUM>, or coupled with the aircraft <NUM> in any other manner. In addition, the propulsor assembly may also be configured in an underwing-mounted configuration (or in any other practical configuration) and provide thrust in response to receiving electrical power from the power generation assembly (and/or an energy storage unit <NUM>). In some embodiments, at least one of the first power assembly <NUM> or the second power assembly <NUM> can be supported by a wing <NUM>, <NUM> of the aircraft <NUM> and at least one of the first power assembly <NUM> or the second power assembly <NUM> can be supported by a fuselage <NUM> of the aircraft <NUM>.

Referring generally to <FIG>, in some embodiments, the hybrid propulsion system <NUM> can generally be configured such that the first power assembly <NUM> has a turbomachine (and, possibly, a prime propulsor (which, for the embodiment of <FIG> are configured together as a gas turbine engine, or rather as a turbofan engine <NUM>)), a first electric machine <NUM> (which for the embodiment depicted in <FIG> is an electric motor/generator) drivingly coupled to the turbomachine, the second power assembly <NUM> (which, for the embodiment of <FIG> is configured as an internal combustion engine <NUM>), a second electric machine <NUM> (which for the embodiment depicted in <FIG> is an electric motor/generator) drivingly coupled to the internal combustion engine <NUM>, the propulsor assembly <NUM> (which for the embodiment of <FIG> is configured as an electric motor assembly), an electric energy storage unit <NUM> (electrically connectable to the first electric machine <NUM>, the second electric machine <NUM>, and/or the propulsor assembly <NUM>), a controller <NUM>, and a power bus <NUM>. The propulsor assembly <NUM>, the electric energy storage unit <NUM>, the first electric machine <NUM>, and the second electric machine <NUM> are each electrically connectable to one another through one or more electric lines <NUM> of the power bus <NUM>. For example, the power bus <NUM> may include various switches or other power electronics <NUM> movable to selectively electrically connect the various components of the hybrid propulsion system <NUM>. Additionally, the power bus <NUM> may further include power electronics <NUM>, such as inverters, converters, rectifiers, etc., for conditioning or converting electrical power within the hybrid propulsion system <NUM>. In some instances, at least one of the first power assembly <NUM> or the second power assembly <NUM> is further configured to provide thrust for the aircraft <NUM> when the at least one of the first power assembly <NUM> or the second power assembly <NUM> is in operation.

As will be appreciated, the controller <NUM> may be configured to distribute electrical power between the various components of the hybrid propulsion system <NUM>. For example, the controller <NUM> may be operable with the power bus <NUM> (including the one or more switches or other power electronics <NUM>) to provide electrical power to or draw electrical power from, the various components, such as the first electric machine <NUM>, the second electric machine <NUM>, and/or the motor of the propulsor assembly <NUM>, to operate the hybrid propulsion system <NUM> between various conditions and perform various functions. Such is depicted schematically as the electric lines <NUM> of the power bus <NUM> extending through the controller <NUM>. In some instances, the controller <NUM> can be configured to provide a desired thrust output from the propulsion assembly <NUM>. In some instances, electrical power is provided from the first power assembly <NUM> to the propulsion assembly <NUM> in response to receiving a command to operate in a first operating condition (such as to accelerate or climb the aircraft <NUM>) and from the second power assembly <NUM> to the propulsion assembly <NUM> in response to receiving a command to operate in a second condition (such as to operate in a level flight condition, an idle condition, or a loiter condition). In some instances, the first power assembly <NUM> may be deactivated when the aircraft <NUM> is in the second operating condition and the second power assembly <NUM> may be deactivated when the aircraft <NUM> is in the first operating condition. Additionally, or alternatively, one of the first and second power assemblies may be utilized to provide a desired amount of thrust for the aircraft <NUM> while the other of the first and second power assembly <NUM> may simultaneously be used for generating electric power for one or more power loads <NUM> (<FIG>) of the aircraft <NUM>. Additionally, or alternatively, the controller <NUM> can be configured to receive a desired thrust output, and if the desired thrust output is within a second power assembly operating range, activate the second power assembly <NUM>, and if the desired thrust output is greater than the second power assembly operating range, activate the first power assembly <NUM>. In some instances, both of the first and second power assemblies <NUM>, <NUM> may generate electrical power simultaneously that can be used by the propulsion assembly <NUM>.

The controller <NUM> may be a stand-alone controller, dedicated to the hybrid propulsion system <NUM>, or alternatively, may be incorporated into one or more of a main system controller for the aircraft <NUM>, a separate controller for the turbofan engine <NUM> (such as a full authority digital engine control system for the turbofan engine <NUM>, also referred to as a FADEC), a separate controller for the internal combustion engine <NUM> (such as a FADEC), etc. For example, the controller <NUM> may be configured in substantially the same manner as the computing system <NUM> described below with reference to <FIG> (and may be configured to perform one or more of the functions of the method <NUM>, described below).

The electric energy storage unit <NUM> may be configured as one or more batteries, one or more capacitors, or any other suitable electrical energy storage devices. It will be appreciated that for the hybrid propulsion system <NUM> described herein, the electric energy storage unit <NUM> is configured to store a relatively large amount of electrical power. For example, in various embodiments, the electric energy storage unit may be configured to store at least about fifty kilowatt-hours of electrical power, such as at least about sixty-five kilowatt-hours of electrical power, such as at least about seventy-five kilowatts hours of electrical power, and up to about one thousand kilowatt-hours of electrical power.

Referring now to <FIG> and <FIG>, the first power assembly <NUM> includes a gas turbofan engine <NUM> mounted, or configured to be mounted, to the first wing <NUM> or the second wing of the aircraft <NUM>. In some embodiments, such as the one illustrated in <FIG>, the gas turbine engine includes a turbomachine <NUM> and a propulsor, the propulsor being a fan (referred to as "fan <NUM>" with reference to <FIG>). Accordingly, for the embodiment of <FIG>, the gas turbine engine is configured as a turbofan engine <NUM>.

The turbofan engine <NUM> defines an axial direction A1 (extending parallel to a longitudinal axis <NUM> provided for reference) and a radial direction R1. As stated, the turbofan engine <NUM> includes the fan <NUM> and the turbomachine <NUM> disposed downstream from the fan <NUM>.

The turbomachine <NUM> depicted generally includes a substantially tubular outer casing <NUM> that defines an annular inlet <NUM>. The outer casing <NUM> encases, in serial flow relationship, a compressor section including a booster or low pressure (LP) compressor <NUM> and a high pressure (HP) compressor <NUM>; a combustion section <NUM>; a turbine section including a first, high pressure (HP) turbine <NUM> and a second, low pressure (LP) turbine <NUM>; and ajet exhaust nozzle section <NUM>. The compressor section, combustion section <NUM>, and turbine section together define at least in part an air flowpath <NUM> through the turbomachine <NUM>.

The turbomachine <NUM> of the turbofan engine <NUM> additionally includes one or more shafts rotatable with at least a portion of the turbine section and, for the embodiment depicted, at least a portion of the compressor section. More particularly, for the embodiment depicted, the turbofan engine <NUM> includes a high pressure (HP) shaft or spool <NUM>, which drivingly connects the HP turbine <NUM> to the HP compressor <NUM>. Additionally, the turbofan engine <NUM> includes a low pressure (LP) shaft or spool <NUM>, which drivingly connects the LP turbine <NUM> to the LP compressor <NUM>.

Further, the fan <NUM> depicted is configured as a variable pitch fan having a plurality of fan blades <NUM> coupled to a disk <NUM> in a spaced apart manner. The fan blades <NUM> extend outwardly from disk <NUM> generally along the radial direction R1. Each fan blade <NUM> is rotatable relative to the disk <NUM> about a respective pitch axis P1 by virtue of the fan blades <NUM> being operatively coupled to a suitable actuation member <NUM> configured to collectively vary the pitch of the fan blades <NUM>. The fan <NUM> is mechanically coupled to the LP shaft <NUM>, such that the fan <NUM> is mechanically driven by the second, LP turbine <NUM>. More particularly, the fan <NUM>, including the fan blades <NUM>, disk <NUM>, and actuation member <NUM>, is mechanically coupled to the LP shaft <NUM> through a power gearbox <NUM>, and is rotatable about the longitudinal axis <NUM> by the LP shaft <NUM> across the power gearbox <NUM>. The power gearbox <NUM> includes a plurality of gears for stepping down the rotational speed of the LP shaft <NUM> to a more efficient rotational fan speed. Accordingly, the fan <NUM> is powered by an LP system (including the LP turbine <NUM>) of the turbomachine <NUM>.

Referring still to the embodiment of <FIG>, the disk <NUM> is covered by rotatable front hub <NUM> aerodynamically contoured to promote an airflow through the plurality of fan blades <NUM>. Additionally, the turbofan engine <NUM> includes an annular fan casing or outer nacelle <NUM> that circumferentially surrounds the fan <NUM> and/or at least a portion of the turbomachine <NUM>. Accordingly, the turbofan engine <NUM> depicted may be referred to as a "ducted" turbofan engine. Moreover, the nacelle <NUM> is supported relative to the turbomachine <NUM> by a plurality of circumferentially-spaced outlet guide vanes <NUM>. A downstream section <NUM> of the nacelle <NUM> extends over an outer portion of the turbomachine <NUM> so as to define a bypass airflow passage <NUM> therebetween.

With further reference to <FIG>, the hybrid propulsion system <NUM> additionally includes the first electric machine <NUM>, which for the embodiment depicted is configured as an electric motor/generator. The first electric machine <NUM> is, for the embodiment depicted, positioned within the turbomachine <NUM> of the turbofan engine <NUM>, inward of the air flowpath <NUM>, and can be coupled to/in mechanical communication with one of the shafts of the turbofan engine <NUM>. For instance, for the embodiment depicted, the electric machine is coupled to the second, LP turbine <NUM> through the LP shaft <NUM>. The first electric machine <NUM> may be configured to convert mechanical power of the LP shaft <NUM> to electrical power (such that the LP shaft <NUM> drives the first electric machine <NUM>), or alternatively, the first electric machine <NUM> may be configured to convert electrical power provided thereto into mechanical power for the LP shaft <NUM> (such that the first electric machine <NUM> drives, or assists with driving, the LP shaft <NUM>). Accordingly, the turbomachine <NUM> may be used to generate electrical power that may be utilized by the propulsion assembly <NUM>, thrust for the aircraft <NUM>, or both.

It will be appreciated that in other embodiments, the first electric machine <NUM> may instead be positioned at any other suitable location within the turbomachine <NUM> or elsewhere. For example, the first electric machine <NUM> may be, in other embodiments, mounted coaxially with the LP shaft <NUM> within the turbine section, or alternatively, may be offset from the LP shaft <NUM> and driven through a suitable gear train. Additionally, or alternatively, in other embodiments, the first electric machine <NUM> may instead be powered by the HP system, e.g., by the HP turbine <NUM> through, e.g., the HP shaft <NUM>, or by both the LP system (e.g., the LP shaft <NUM>) and the HP system (e.g., the HP shaft <NUM>) via a dual drive system. Additionally, or alternatively, still, in other embodiments, the first electric machine <NUM> may include a plurality of electric machines, e.g., with one being drivingly connected to the LP system (e.g., the LP shaft <NUM>) and one being drivingly connected to the HP system (e.g., the HP shaft <NUM>). Further, although the first electric machine <NUM> is described as an electric motor/generator, in other embodiments, the first electric machine <NUM> may be configured solely as an electric generator.

In various embodiments, the first electric machine <NUM> may be configured to generate at least about ten kilowatts of electrical power when driven by the turbomachine <NUM>, such as at least about fifty kilowatts of electrical power, such as at least about sixty-five kilowatts of electrical power, such as at least about seventy-five kilowatts of electrical power, such as at least about one hundred kilowatts of electrical power, such as up to five thousand kilowatts of electrical power, such as up to eight hundred and fifty kilowatts of electrical power, such as up to nine hundred kilowatts of electrical power. Additionally, or alternatively, the first electric machine <NUM> may be configured to provide, or otherwise add, horsepower (hp) of mechanical power to the turbomachine <NUM> when the first electric machine <NUM> is provided electrical power from, e.g., the electric energy storage unit <NUM> of the second power assembly <NUM>. For example, in various embodiments, the first electric machine <NUM> may be configured to provide at least about fifteen, such as at least about fifty horsepower of mechanical power to the turbomachine <NUM>, such as at least about seventy-five horsepower, such as at least about one hundred horsepower, such as at least about one hundred and twenty horsepower, such as up to about seven thousand horsepower.

Referring still to <FIG> and <FIG>, the turbofan engine <NUM> further includes a controller <NUM>, such as a FADEC, and a plurality of sensors. The controller <NUM> of the turbofan engine <NUM> may be configured to control operation of, e.g., the actuation member <NUM>, the fuel delivery system, etc. Additionally, referring back also to <FIG>, the controller <NUM> of the turbofan engine <NUM> is operably connected to the controller <NUM> of the hybrid propulsion system <NUM>. Moreover, the controller <NUM> may further be operably connected to one or more of the first power assembly <NUM> (including controller <NUM>), the first electric machine <NUM>, the second power assembly <NUM>, the second electric machine <NUM>, the propulsor assembly <NUM>, and the energy storage unit <NUM> through a suitable wired or wireless communication system (depicted in phantom).

In various embodiments, the turbofan engine <NUM> may further include one or more sensors positioned to, and configured to, sense data indicative of one or more operational parameters of the turbofan engine <NUM>. For example, the turbofan engine <NUM> may include one or more temperature sensors configured to sense a temperature within a air flowpath <NUM> of the turbomachine <NUM>. For example, such sensors may be configured to sense an exhaust gas temperature at an exit of the combustion section <NUM>. Additionally, or alternatively, the turbofan engine <NUM> may include one or more pressure sensors to sense data indicative of a pressure within the air flowpath <NUM> of the turbomachine <NUM>, such as within a combustor within the combustion section <NUM> of the turbomachine <NUM>. Further, in still other embodiments, the turbofan engine <NUM> may also include one or more speed sensors configured to sense data indicative of a rotational speed of one or more components of the turbofan engine <NUM>, such as one or more of the LP spool <NUM> or the HP spool <NUM>. Additionally, in various embodiments, the turbofan engine <NUM>, the hybrid propulsion system <NUM> as a whole, and/or an aircraft <NUM> incorporating the hybrid propulsion system <NUM>, may include one or more ambient conditions sensors, such as one or more ambient temperature sensors, positioned outside the air flowpath <NUM> of the turbomachine <NUM> for sensing data indicative of an ambient condition, such as an ambient temperature. Accordingly, in at least various embodiments, the hybrid propulsion system <NUM> may receive information regarding one or more ambient conditions from the aircraft <NUM>. Notably, however, in other embodiments, ambient conditions may be sensed within the air flowpath <NUM> of the turbomachine <NUM>, e.g., at the inlet <NUM>.

It should further be appreciated that the turbofan engine <NUM> depicted in <FIG> may, in other embodiments, have any other suitable configuration. For example, in other embodiments, the fan <NUM> may not be a variable pitch fan, and further, in other embodiments, the LP shaft <NUM> may be directly mechanically coupled to the fan <NUM> (e.g., the turbofan engine <NUM> may not include the gearbox <NUM>). Further, it should be appreciated that in other embodiments, the turbofan engine <NUM> may be configured as any other suitable gas turbine engine. For example, in other embodiments, the turbofan engine <NUM> may instead be configured as a turboprop engine, an unducted turbofan engine, a turbojet engine, a turboshaft engine, etc..

Referring now to <FIG> and <FIG>, as previously stated, the hybrid propulsion system <NUM> can additionally include the second power assembly <NUM> mounted, for the embodiment depicted in <FIG>, to the fuselage <NUM> (directly or indirectly) and/or one or more of the first and the second wings <NUM>, <NUM> of the aircraft <NUM>. As illustrated in <FIG>, in some embodiments, the second power assembly <NUM> can be generally configured as an internal combustion engine <NUM> and a propeller. The internal combustion engine <NUM> defines an axial direction A<NUM> (extending parallel to a longitudinal centerline <NUM> provided for reference) and a radial direction R<NUM>.

In some embodiments, the internal combustion engine <NUM> includes an engine block <NUM> that forms the main structure of the internal combustion engine <NUM> and contains and/or defines many of the internal features of the internal combustion engine <NUM>. The engine block <NUM> is constructed and arranged to define a crankcase <NUM> and a plurality of cylinders <NUM>. In the various embodiments, the crankcase <NUM> is oriented substantially parallel to a longitudinal centerline <NUM> of the internal combustion engine <NUM>. The crankcase <NUM> houses a crankshaft <NUM> that is disposed along the longitudinal centerline <NUM>.

Referring still to <FIG> and <FIG>, the plurality of cylinders <NUM> can include two to twelve (or more) cylinders, such as four to twelve cylinders, such as four to eight cylinders, and such as six cylinders. The cylinders <NUM> are arranged so that they extend upward from the crankcase <NUM>. Each cylinder <NUM> can extend at an angle relative to a radial direction R2 that is perpendicular to the longitudinal centerline. As the number of cylinders <NUM> is increased, for example to six cylinders, the cylinders <NUM> can be alternated on opposite sides of the radial direction R2 in a configuration that may be referred to in the art as a "V" configuration, thereby creating a "V-type" internal combustion engine <NUM> with three cylinders on each side of the internal combustion engine <NUM>. It is understood that two cylinders may be substantially opposed to one another, rather than a full alternated arrangement, to save space.

In various embodiments, each cylinder <NUM> is constructed to slidably receive a piston <NUM> that is operatively connected to the crankshaft <NUM> via a connecting rod <NUM>. Each connecting rod <NUM> is rotatably connected to one of the pistons <NUM> at one end portion and rotatably connected to the crankshaft <NUM> via a pin-type crankshaft journal <NUM> at the opposite end portion. The pistons <NUM> reciprocate linearly within the cylinders <NUM>. In turn, the connecting rods <NUM> convert the linear movement of the pistons <NUM> into rotational movement of the crankshaft <NUM>, and vice-versa.

In some embodiments, the crankcase <NUM> can include at least one crank chamber <NUM>, and in various embodiments, the crankcase <NUM> can include one isolated crank chamber <NUM> for each pair of substantially opposed cylinders <NUM>. A bore <NUM> can extend through the crankcase <NUM> and each of the crank chambers <NUM>. The crankshaft <NUM> is received by the bore <NUM>. In some examples, a balancing shaft <NUM> can also extend through the crankcase <NUM>. The balancing shaft <NUM> is provided to counteract the moment generated by rotation of the crankshaft <NUM> and the piston assembly which produce mass moment unbalancing of the first order. The balancing shaft <NUM> and the crankshaft <NUM> extend through the crankcase <NUM> in a parallel relationship, as shown in <FIG>. The balancing shaft <NUM> is rotatably mounted within a bore <NUM> that extends through the crankcase <NUM>. Suitable bearing assemblies are provided for smooth rotation of the balancing shaft <NUM>. The balancing shaft <NUM> is operatively connected to the crankshaft <NUM> through a gear <NUM>, which may be located within a gearbox <NUM> at one end portion of the crankcase <NUM>.

In some embodiments, an air intake system <NUM> can be constructed and arranged to receive air from the environment and deliver the air to intake passageways. A throttle valve <NUM> can be disposed within an entry of the air intake system <NUM> and can be controlled by a controller <NUM>. The throttle valve <NUM> is mechanically or electrically movable to increase or decrease the amount of air that enters air intake system <NUM>, and thus assists in controlling the speed of rotation of the crankshaft <NUM>. It will be appreciated that, in some embodiments, the internal combustion engine <NUM> can include a turbocharger that can be mounted to the internal combustion engine <NUM>. In such embodiments, the turbocharger can include an internal turbine, which in turn drives a compressor that is used to compress the intake air. Thus, the turbocharger can be designed to increase the pressure of the incoming air to the air intake system <NUM>.

The internal combustion engine <NUM> described herein can be configured to provide a total engine output of about one hundred and forty to about six hundred horsepower (hp). For example, the total engine output is about one hundred and fifty to about five hundred horsepower, about one hundred and sixty to about four hundred horsepower, about one hundred seventy to about three hundred seventy-five horsepower, and/or about one hundred eighty to about three hundred fifty horsepower. In the various embodiments, the total engine output can be about two hundred twenty horsepower for a naturally aspirated internal combustion engine <NUM>, and about three hundred horsepower for a turbocharged internal combustion engine <NUM>. In some embodiments, the second power assembly <NUM> (or the internal combustion engine <NUM>) may be configured to rotate and generate electrical power.

In some embodiments, the internal combustion engine <NUM> may further include a propeller shaft <NUM> that can be operatively connected to the internal combustion engine <NUM>, and is also operatively connected to a propeller <NUM>. For instance, the propeller shaft <NUM> is connected to the propeller <NUM> at one end portion and a gearbox <NUM>, at an opposite end portion. In some embodiments, the gearbox <NUM> is constructed and arranged to rotate the propeller shaft <NUM>, and hence the propeller <NUM>, at a speed of about one hundred to about three thousand revolutions per minute when the internal combustion engine <NUM> is operating under normal conditions.

Further, in embodiments including the propeller <NUM>, a variable pitch fan having a plurality of propeller blades <NUM> may be coupled to a disk <NUM> in a spaced apart manner. The propeller blades <NUM> extend outwardly from the disk <NUM> generally along the radial direction R<NUM>. Each propeller blade <NUM> is rotatable relative to the disk <NUM> about a respective pitch axis P2 by virtue of the propeller blades <NUM> being operatively coupled to a suitable actuation member <NUM> configured to collectively vary the pitch of the propeller blades <NUM>. The propeller <NUM> is mechanically coupled to the propeller shaft <NUM>, such that the propeller <NUM> is mechanically driven by the crankshaft. More particularly, the propeller <NUM>, including the propeller blades <NUM>, the disk <NUM>, and the actuation member <NUM>, is mechanically coupled to the crankshaft <NUM> through the gearbox <NUM>, which can also be referred to as a speed reduction unit or a propeller speed reduction unit. Accordingly, the propeller <NUM> can be powered by a crankshaft <NUM> of the internal combustion engine <NUM>.

In some embodiments, the second electric machine <NUM> may be disposed at one or both end portions of the crankshaft <NUM>. The second electric machine <NUM> may be configured to convert mechanical power of the crankshaft <NUM> to electrical power when the crankshaft <NUM> drives the second electric machine <NUM> through coupling with the crankshaft <NUM> or through an additional shaft <NUM> that rotates with the crankshaft <NUM>. In turn, the electrical power is transmitted from the second electric machine <NUM> to one or more electric lines <NUM>. Alternatively, the second electric machine <NUM> may be configured to convert electrical power provided thereto into mechanical power for the crankshaft <NUM> such that the second electric machine <NUM> drives, or assists with driving, the crankshaft <NUM> through the one or more electric lines <NUM>.

It will be appreciated that in some embodiments, the second electric machine <NUM> may instead be positioned at any other suitable location within the internal combustion engine <NUM> or elsewhere. For example, the second electric machine <NUM> may be, in other embodiments, mounted coaxially with the propeller shaft <NUM> (or coupled with the propeller shaft in lieu of a propeller), or alternatively may be offset from the crankshaft <NUM> and/or the propeller shaft <NUM> and driven through a suitable gear train. Further, although the second electric machine <NUM> is described as an electric motor/generator, in other embodiments, the second electric machine <NUM> may be configured solely as an electric generator.

Notably, in various embodiments, the second electric machine <NUM> may be configured to generate at least about ten kilowatts of electrical power when driven by the internal combustion engine <NUM>, such as at least about fifty kilowatts of electrical power, such as at least about sixty-five kilowatts of electrical power, such as at least about seventy-five kilowatts of electrical power, such as at least about one hundred kilowatts of electrical power, such as up to four hundred kilowatts of electrical power.

Referring still to <FIG> and <FIG>, the internal combustion engine <NUM> further includes a controller <NUM>, which may be a FADEC system, and a plurality of sensors. The controller <NUM> of the internal combustion engine <NUM> may be configured to monitor and control many of the operating parameters of the internal combustion engine <NUM>. For example, the controller <NUM> can monitor and control the air-to-fuel ("air/fuel") ratio, or fuel richness, that is provided to the combustion chambers. This can be done by controlling the amount of fuel that is injected. The controller <NUM> can additionally or alternatively monitor and control the rotational speed of the crankshaft <NUM>, which can be done by controlling the amount of fuel and air that is provided to the combustion chambers. The controller <NUM> can also provide propeller pitch control, which allows the internal combustion engine <NUM> operate more efficiently. Additionally, referring back also to <FIG>, the controller <NUM> of the internal combustion engine <NUM> is operably connected to the controller <NUM> of the hybrid propulsion system <NUM>.

Referring now particularly to <FIG> and <FIG>, as previously stated the hybrid propulsion system <NUM> can additionally include the propulsor assembly <NUM> mounted, for the embodiment depicted in <FIG>, to the first and second wings <NUM>, <NUM> of the aircraft <NUM>. As illustrated in <FIG>, in some embodiments, the propulsor assembly <NUM> can include an electric motor <NUM> and at least one aerodynamic propulsor/fan <NUM>/<NUM>. The propulsor assembly <NUM> defines an axial direction A3 extending along a longitudinal centerline axis <NUM> that extends therethrough for reference, as well as a radial direction R3. For the embodiment depicted, the propulsor <NUM> is rotatable about the centerline axis <NUM> by the electric motor <NUM>.

The at least one aerodynamic propulsor <NUM>, in some embodiments, can include a plurality of fan blades <NUM> and a fan shaft <NUM>. The plurality of fan blades <NUM> are attached to/rotatable with the fan shaft <NUM> and spaced generally along a circumferential direction of the propulsor assembly <NUM>. In various embodiments, the plurality of fan blades <NUM> may be attached in a fixed manner to the fan shaft <NUM>, or alternatively, the plurality of fan blades <NUM> may be rotatable relative to the fan shaft <NUM>, such as in the embodiment depicted. For example, the plurality of fan blades <NUM> each define a respective pitch axis P3, and for the embodiment depicted are attached to the fan shaft <NUM> such that a pitch of each of the plurality of fan blades <NUM> may be changed, e.g., in unison, by a pitch change mechanism <NUM>. Changing the pitch of the plurality of fan blades <NUM> may increase an efficiency of the propulsor assembly <NUM> and/or may allow the propulsor assembly <NUM> to achieve a desired thrust profile. With such an embodiment, the propulsor <NUM> may be referred to as a variable pitch fan.

Moreover, for the embodiment depicted, the propulsor assembly <NUM> depicted additionally includes a fan casing or outer nacelle <NUM>, attached to a core <NUM> of the propulsor assembly <NUM> through one or more struts or outlet guide vanes <NUM>. For the embodiment depicted, the outer nacelle <NUM> substantially completely surrounds the at least one aerodynamic propulsor <NUM>, and particularly the plurality of fan blades <NUM>. Accordingly, for the embodiment depicted, the propulsor assembly <NUM> may be referred to as a ducted electric fan. In some embodiments, however, the propulsor assembly <NUM> may be configured as an unducted fan that does not include the outer nacelle <NUM>.

Referring still to <FIG>, the fan shaft <NUM> is mechanically coupled to the electric motor <NUM> within the core <NUM>, such that the electric motor <NUM> drives the propulsor <NUM> through the fan shaft <NUM>. The fan shaft <NUM> is supported by one or more bearings <NUM>, such as one or more roller bearings, ball bearings, or any other suitable bearings. Additionally, the electric motor <NUM> may be an inrunner electric motor (e.g., including a rotor positioned radially inward of a stator), or alternatively may be an outrunner electric motor (e.g., including a stator positioned radially inward of a rotor), or alternatively, still, may be an axial flux electric motor (e.g., with the rotor neither outside the stator nor inside the stator, but rather offset from it along the axis of the electric motor).

As briefly noted above, the electrical power source (e.g., the first electric machine <NUM>, the second electric machine <NUM>, and/or the electric energy storage unit <NUM>) is electrically connected with the propulsor assembly <NUM> (e.g., the electric motor <NUM>) for providing electrical power to the propulsor assembly <NUM>. For instance, in some embodiments, the electric motor <NUM> is in electrical communication with the first electric machine <NUM>, the second electric machine <NUM>, and/or the electric energy storage unit <NUM> through the electrical power bus <NUM>, and more particularly, through the one or more electrical cables or lines <NUM> extending therebetween.

Referring still to <FIG> and <FIG>, the propulsor assembly <NUM> further includes a controller <NUM>, which may be a FADEC system, and a plurality of sensors. The controller <NUM> of propulsor assembly <NUM> may be configured to monitor and control many of the operating parameters of the electric motor <NUM>. For example, the controller <NUM> can monitor and control the rotational speed of the at least one aerodynamic propulsor, the pitch of the at least one aerodynamic propulsor, etc. Additionally, referring back also to <FIG>, the controller <NUM> of the propulsor assembly <NUM> is operably connected to the controller <NUM> of the hybrid propulsion system <NUM>.

It should be appreciated that in various embodiments the hybrid propulsion system <NUM> may have any other suitable configuration, and further, may be integrated into an aircraft <NUM> in any other suitable manner. For example, in other embodiments, the propulsor assembly <NUM> of the hybrid propulsion system <NUM> may instead be configured as a plurality of propulsor assemblies <NUM> that are distributed about the aircraft in any practicable manner while being electrically coupled with the first power assembly <NUM>, the second power assembly <NUM>, the energy storage unit <NUM>, or any combination thereof.

In various embodiments, the electric propulsor assembly(ies) <NUM>, the gas turbine engine(s), the first electric machine(s) <NUM>, the internal combustion engine(s) <NUM>, and the second electric machines <NUM> may be mounted to the aircraft <NUM> at any other suitable location in any other suitable manner (including, e.g., tail mounted configurations). For example, in various embodiments, the electric propulsor assembly may be configured to ingest boundary layer air and reenergize such boundary layer air to provide a propulsive benefit for the aircraft <NUM> (the propulsive benefit may be thrust, or may simply be an increase in overall net thrust for the aircraft <NUM> by reducing a drag on the aircraft <NUM>).

Referring to <FIG> and <FIG>, the hybrid propulsion system <NUM> includes the first electric machine <NUM>, the second electric machine <NUM>, and the electric energy storage unit <NUM> electrically connectable to the electric motor <NUM> of the propulsion assembly <NUM>. It will be appreciated that the hybrid propulsion system <NUM> provided herein may include any number of first power assemblies, second power assemblies, and/or propulsor assemblies located on any portion of the aircraft <NUM>.

The first electric machine <NUM> is additionally coupled to the turbomachine <NUM>. In such a manner, the first electric machine <NUM> may extract power from the turbomachine <NUM> and/or provide power to the first turbomachine <NUM>. The second electric machine <NUM> is additionally coupled to the internal combustion engine <NUM>. In such a manner, the second electric machine <NUM> may extract power from the internal combustion engine <NUM> and/or provide power to the internal combustion engine <NUM>.

With further reference to <FIG>, in some embodiments, the first power assembly <NUM> can be configured as a continuous combustion engine, such as a turbine engine, turboprop engine, an unducted turbofan engine <NUM>, a turbojet engine, a turboshaft engine, etc. In various embodiments, the first power assembly <NUM> may be capable of producing up to three thousand horsepower, such as five hundred horsepower to two thousand five hundred horsepower, such as one thousand horsepower to two thousand horsepower, such as one thousand eight hundred horsepower to two thousand five hundred horsepower.

The first electric machine <NUM> may be integrated within the first power assembly <NUM> and configured as motor/generator. In various embodiments, the first electric machine <NUM> may be configured to generate a first amount of electrical power, which may be up to one megawatt (mW) during operation of the first power assembly <NUM>, such as up to nine hundred kilowatts (kW), such as up to eight hundred fifty kW when operated in a generator mode.

In some embodiments, the second power assembly <NUM> may be configured as a reciprocating engine and/or a fuel cell assembly that is coupled to the propulsion assembly <NUM>. For example, the second power assembly <NUM> may be configured as an internal combustion engine <NUM> that can be configured as a two-stroke engine (e.g., clerk cycle, day cycle, etc.), a Four-stroke engine (e.g., Otto cycle), a six-stroke engine, or any other number of strokes. In addition, the internal combustion engine <NUM> may be configured as compression-ignition engine and/or a spark-ignition engine. Further, the internal combustion engine <NUM> may be configured to operate through a mechanical/thermodynamical cycle (e.g., Atkinson cycle, Miller cycle, etc.) and/or a rotary engine (e.g., a Wankel engine). In various embodiments, the second power assembly <NUM> may be configured to produce up to seven hundred fifty horsepower, such as one hundred to seven hundred horsepower, such as two hundred to five hundred horsepower.

The second electric machine <NUM> may be operably coupled to the second power assembly <NUM> and configured as motor/generator. In various embodiments, the second electric machine <NUM> may be configured to generate a second amount of electrical power, which may be up to eight hundred kW during operation of the first power assembly <NUM>, such as between fifty and five hundred kW, such as between one hundred and four hundred fifty kW, such as between one hundred fifty and three hundred and seventy five kW, or any other practicable range when operated in a generator mode. In some embodiments, the second electric machine <NUM> may be configured to generate less electrical power than the first electric machine <NUM>. For instance, the second electric machine <NUM> may generate less than half of the electrical power of the first electric machine <NUM>.

In some embodiments, the first power assembly <NUM> is drivingly coupled to the first electric machine <NUM> to produce a first amount of electric power and the second power assembly <NUM> is drivingly coupled to the second electric machine <NUM> to produce a second amount of electric power. In some instances, the second power assembly <NUM> can be configured to generate the second amount of electric power more efficiently than the first power assembly <NUM> generates the first amount of electric power. For instance, as provided herein, the first power assembly <NUM> may be a continuous combustion engine <NUM> (or any other type of engine) that operates at a first efficiency at an idle speed (e.g. <NUM>-<NUM>% of maximum output), a second efficiency at an mid-range operating speed (e.g. <NUM>-<NUM>% of maximum output), and/or a third efficiency at a maximum output (e.g. <NUM>-<NUM>% of maximum output). The second power assembly <NUM> may be an internal combustion engine <NUM> (or any other type of engine) that operates at a fourth efficiency at an idle speed (e.g. <NUM>-<NUM>% of maximum output), a fifth efficiency at an mid-range operating speed (e.g. <NUM>-<NUM>% of maximum output), and/or a sixth efficiency at a maximum output (e.g. <NUM>-<NUM>% of maximum output). In some instances, each of the first and second power assemblies <NUM>, <NUM> may use a combustible fuel in order to operate at the defined efficiencies.

In some instances, the first power assembly <NUM>, when configured as a turbomachine, is most efficient at maximum power output making the third efficiency greater than the first efficiency. Conversely, when operated at lower rotational speeds, the pressure of the compressed air within the first power assembly <NUM> drops and thus thermal and fuel efficiency drop dramatically within the first power assembly <NUM>. Accordingly, the efficiency of the first power assembly <NUM> can steadily decline with reduced power output and can be lower in the low power range. Conversely, the fourth, the fifth and the sixth efficiencies may be generally within a predefined percentage (such as <NUM>%) of each other. In some instances, the fourth, the fifth and the sixth efficiencies may be greater than that of the first efficiency and/or less than that of the third efficiency. It is to be understood that the efficiencies provided herein are related to the amount of fuel consumed to generate an amount of propulsion through the propulsor assembly.

Additionally or alternatively, in various embodiments, the first power assembly first power assembly <NUM> can be configured to operate in a first range of revolutions per minute. The second power assembly <NUM> is configured to operate in a second range of revolutions per minute that is at least partially less than or greater than the first range of revolutions per minute. In some instances, the variation in revolutions per minute between the first and second power assemblies <NUM>, <NUM> may be fairly large due to various engine designs between the first and second power assemblies.

Additionally or alternatively, in various embodiments, the first power assembly generates a first noise level while operating at an idle speed, or at any other defined speed, while the second power assembly generates a second noise level while operating at an idle speed, or at any other defined speed. In various embodiments, the first noise level may be greater than the second noise level while operating at a common operating speed and/or operating parameter.

In some embodiments, the propulsor assembly <NUM> can include one or more electric motors and it will be appreciated that for the embodiment depicted in <FIG>, the propulsor assembly <NUM> can be configured as a pure electric propulsor assembly in which the electric motor <NUM> of the propulsor assembly <NUM> is coupled independently to a propulsor <NUM>. In other embodiments, such as the one depicted in <FIG>, the propulsor assembly <NUM> can be configured as part of a hybrid propulsor in which at least one of the first and/or second power assemblies may turn a common propulsor <NUM> with at least one electrical motor <NUM> of the propulsor assembly <NUM>. In various embodiments, the power loading of each electrical motor <NUM> within the propulsor assembly <NUM> may be up to <NUM><NUM>/m (twenty pound feet/horsepower (lbf/hp)), such as between <NUM> to <NUM><NUM>/m (one to ten pound feet /horsepower), such as between <NUM> and <NUM><NUM>/m (two and eight pound feet/horsepower). In some embodiments, the propulsor assembly <NUM> may include more than one motor and/or more than one propulsor. Each of the propulsors <NUM> may be ducted and/or unducted and distributed about the aircraft <NUM> in any manner. In various embodiments, the propulsor assembly <NUM> may include any number of motor(s) that may be operably coupled to any one or more propulsors <NUM>. Each of the one or more propulsors <NUM> may be operably coupled in parallel and/or in series.

In some embodiments, The first and the second power assemblies <NUM>, <NUM> are powered in parallel. Additionally or alternatively, the first and the second power assemblies <NUM>, <NUM> are powered in parallel. Accordingly, the disparate first and the second power assemblies <NUM>, <NUM> may be coupled in series electric transfer as well as parallel for transferring power based on a mechanical coupling of the first and the second power assemblies <NUM>, <NUM> to one another. For example, the first power assembly <NUM> may provide power to the power electronics <NUM> through the first electric machine <NUM> and/or through rotation of the second electric machine <NUM>. Likewise, the second power assembly <NUM> may provide power to the power electronics <NUM> through the second electric machine <NUM> and/or through rotation of the first electric machine <NUM> that is operably coupled with the first power assembly <NUM>.

As is also depicted in <FIG>, the hybrid propulsion system <NUM> further includes a controller <NUM> and a power bus <NUM>. Various components of the first power assembly <NUM>, the second power assembly <NUM>, the propulsor assembly <NUM>, and the electric energy storage unit <NUM> are each electrically connectable to one another through one or more electric lines <NUM> of the power bus <NUM>. For example, the power bus <NUM> may include various switches or other power electronics <NUM> movable to selectively electrically connect the various components of the hybrid propulsion system <NUM>, and optionally to convert or condition such electrical power transferred therethrough. Accordingly, in certain operations, the first electric machine <NUM> may provide electrical power to the propulsor assembly <NUM>, or vice versa. Further, in certain operations, the first electric machine <NUM> may provide electrical power to the second electric machine <NUM> and/or the second power assembly <NUM>, or vice versa. Likewise, in certain operations, the second power assembly <NUM> and/or the second electric machine <NUM> may provide electrical power to the propulsor assembly <NUM>, or vice versa. Additionally, or alternatively, the first electric machine <NUM> and/or the second electric machine <NUM> may provide electrical power to the electric energy storage unit <NUM>, or the electric energy storage unit <NUM> may provide electrical power to the first power assembly <NUM>, the second power assembly <NUM>, and/or the propulsor assembly <NUM>. In various embodiments various configurations of thrust outputs may occur when each power assembly <NUM>, <NUM> is switched on or off, based on the system <NUM> power demand, and each power assembly <NUM>, <NUM> can be throttled back while another power assembly <NUM>, <NUM> is brought up to speed during the power transition.

With reference to <FIG>, the hybrid propulsion system <NUM> can include one or more propulsors <NUM> that are operably coupled to more than one of the first power assembly <NUM>, the second power assembly <NUM>, and/or the propulsor assembly <NUM>. For example, in some embodiments, a propulsor <NUM> may be operably coupled with the first power assembly <NUM> and the propulsor assembly <NUM>. In some embodiments, the propulsor <NUM> may additionally be operably coupled with the second power assembly <NUM>. Alternatively, in some embodiments, the propulsor <NUM> may be coupled with the second power assembly <NUM> and the propulsor assembly <NUM>. Additionally, or alternatively, a first propulsor <NUM> may be coupled with the first power assembly <NUM> and the propulsor assembly <NUM> while a second propulsor <NUM> is operably coupled with the second power assembly <NUM> and the propulsor assembly <NUM>.

One or more variable thrust assemblies <NUM> may be positioned within the hybrid propulsion system <NUM> to allow for selective coupling of the at least one aerodynamic propulsor(s) <NUM> to one or more of the power assemblies <NUM>, <NUM> and the propulsion assembly(ies) <NUM>. For instance, in various embodiments, the variable thrust assemblies <NUM> may be configured as a clutch assembly and/or a pitch change mechanism <NUM> (<FIG>). In some embodiments, the propulsor assembly <NUM> may rotate the propulsor <NUM> through the variable thrust assembly <NUM> while the first and/or second power assemblies <NUM>, <NUM> are disengaged from the fan. While disengaged, the first and/or second power assemblies <NUM>, <NUM> may continue to generate electrical power through the respective first and second electric machines <NUM>, <NUM>. The generated electrical power may be used by the propulsor assembly <NUM>, provided to the energy storage unit <NUM>, and/or used by one or more power loads <NUM> of the aircraft <NUM>.

As depicted in <FIG>, each of the first power assembly <NUM>, the second power assembly <NUM>, and the propulsor assembly <NUM> may be operably coupled to the controller <NUM> and one or more common power electronics <NUM>. In some instances, the power electronics <NUM> can be configured, for example, to provide or enable power conversion operations (e.g. AC to DC conversion, DC to AC conversion, a first DC power to a second DC power, etc.) to selectively enable or disable the delivery of power to one or more particular propulsor assemblies and/or power loads <NUM>, depending on, for example, available power distribution supply, criticality of electrical load functionality, or aircraft mode of operation, such as take-off, cruise, loiter, or ground operations.

Referring now to <FIG>, a flow diagram of a method <NUM> for operating a hybrid propulsion system <NUM> of an aircraft <NUM> is provided. The method <NUM> may generally be operable with one or more of the hybrid propulsion systems <NUM> described above with reference to <FIG>. For example, the hybrid propulsion system <NUM> may generally include a first power assembly <NUM> operably coupled with a first electric machine <NUM>, a second power assembly <NUM> operably coupled with a second electric machine <NUM>, and a propulsor assembly <NUM> including a propulsor <NUM>. In some instances, the propulsor assembly <NUM> can be operably coupled to at least one of the first electric machine <NUM> and the second electric machine <NUM>.

As is depicted, the method <NUM> includes at (<NUM>) receiving, by one or more computing devices, a command to provide a desired thrust output. In some instances, receiving, by the one or more computing devices, the command to provide the desired thrust output at (<NUM>) includes at (<NUM>) receiving, by one or more computing devices, a command to provide a first amount of thrust output for a first operating condition. In various instances, the first operating condition may include a pre-level flight condition, such as a takeoff flight condition or a climb flight condition. In such instances, at (<NUM>), the method includes providing, by the one or more computing devices, a first amount of electrical power from the first electric machine <NUM> to provide the first thrust output from the first power assembly <NUM> to the propulsor assembly <NUM>.

In some instances, receiving, by the one or more computing devices, the command to provide the desired thrust output at (<NUM>) includes at (<NUM>) receiving, by the one or more computing devices, a command to provide a second amount of thrust output for a second operating condition. In various embodiments, the second operating condition can be indicated by a command to operate the aircraft <NUM> in a cruise condition, an idle condition, a loiter condition. In such instances, at (<NUM>), the method includes providing, by the one or more computing devices, a second amount of electrical power from the second electric machine <NUM> to provide the second thrust output from the second power assembly <NUM> to the propulsor assembly <NUM>.

In some instances, if the desired thrust output is within a second power assembly operating range, which may be indicative of a cruise condition and/or a loiter condition, the second power assembly <NUM> may be activated. For instance, in some embodiments, the internal combustion engine <NUM> of the second power assembly <NUM> may operate in a defined range of revolutions per minute (rpm) while functioning in the second power assembly operating range. Further, the second power assembly operating range may alternatively be defined as any operating parameter of the internal combustion engine <NUM> that can monitor the operation thereof and generally ensure that the internal combustion engine <NUM> is operating in a manner that is consistent with operating the internal combustion engine <NUM> in a defined range for that parameter. Moreover, in some instances, the first amount of electrical power can be greater than a maximum electrical power output of the second power assembly <NUM>.

If the desired thrust output is greater than the second power assembly operating range, which may be indicative of a pre-cruise flight condition, including a takeoff flight condition, a climb flight condition, or any other high-speed conditions, the first power assembly <NUM> may be activated. In some embodiments, the first power assembly <NUM> may be deactivated when the aircraft <NUM> is in the second operating condition or when the desired thrust output is within the second power assembly operating range.

The method <NUM> can further includes at (<NUM>) receiving, by the one or more computing devices, data indicative of a parameter approaching or exceeding an upper threshold. Notably, as used herein, the term "approaching or exceeding" refers to a parameter value being within a predetermined range of a threshold, or being above the threshold. In certain aspects, such as the aspect depicted, receiving, by the one or more computing devices, data indicative of the parameter approaching or exceeding the upper threshold at (<NUM>) includes at (<NUM>) receiving, by the one or more computing devices, data indicative of a temperature approaching or exceeding a threshold for the first power assembly <NUM> or the second power assembly <NUM>. The temperature threshold may be a temperature threshold above which the first power assembly <NUM> is limited in an amount of effective output power it may produce by virtue of the ingested ambient air being too hot. However, in other aspects of the present disclosure, the temperature parameter may be any other suitable temperature parameter. For example, the temperature parameter may include an exhaust gas temperature parameter approaching or exceeding an upper exhaust gas temperature parameter threshold.

Moreover, the aspect depicted further includes at (<NUM>) providing or assist with providing, by the one or more computing devices, the first thrust output in response to receiving the command to operate the aircraft <NUM> in the first operating condition at (<NUM>) and receiving the data indicative of the parameter approaching or exceeding the upper threshold at (<NUM>) through at least one of the second power assembly <NUM> and the energy storage unit <NUM> to generate the first thrust output.

By providing the electrical power to the propulsion assembly <NUM> in accordance with one or more aspects of the present disclosure, the hybrid propulsion system <NUM> may provide the desired thrust output in a more efficient manner from the first and/or second power assembly <NUM>, <NUM> to the propulsor assembly <NUM>.

By operating in accordance with one or more these aspects, the hybrid propulsion system <NUM> provided in <FIG> and/or the method <NUM> provided in <FIG> may provide a sufficient amount of thrust output to the aircraft at each operation mode of the aircraft (e.g., takeoff, cruise, loiter, etc.) while operating in an efficient manner. For instance, the hybrid propulsion system <NUM> provided herein may use a first power assembly <NUM> that incorporates a lightweight, high specific power engine for takeoff and/or dash conditions (e.g., a constant combustion engine) and a second power assembly <NUM> that incorporates a high efficiency, low specific power engine for long-duration cruise/loiter conditions (e.g., an intermittent combustion engine). In addition, the hybrid propulsion system may further incorporate a propulsor assembly <NUM> that uses a partial or full electrical drive-train that can enable the combination of power from disparate plants in order to drive one or more propulsors. In various embodiments, through one or more computing systems, various configurations of thrust outputs may occur when each propulsor assembly is switched on or off, based on the system power demand, and each propulsor assembly can be throttled back while another propulsor assembly is brought up to speed during the power transition.

Referring now to <FIG>, an example computing system <NUM> according to example embodiments of the present disclosure is depicted. The computing system <NUM> can be used, for example, as a controller <NUM> in a hybrid propulsion system <NUM>, the controller <NUM> of the first power assembly <NUM>, the controller <NUM> of the second power assembly <NUM>, and/or the controller <NUM> of the propulsor assembly <NUM>. The computing system <NUM> can include one or more computing device(s) <NUM>. The computing device(s) <NUM> can include one or more processor(s) 510A and one or more memory device(s) 510B. The one or more processor(s) 510A can include any suitable processing device, such as a microprocessor, microcontroller, integrated circuit, logic device, and/or other suitable processing device. The one or more memory device(s) 510B can include one or more computer-readable media, including, but not limited to, non-transitory computer-readable media, RAM, ROM, hard drives, flash drives, and/or other memory devices.

The one or more memory device(s) 510B can store information accessible by the one or more processor(s) 510A, including computer-readable instructions 510C that can be executed by the one or more processor(s) 510A. The instructions 510C can be any set of instructions that when executed by the one or more processor(s) 510A, cause the one or more processor(s) 510A to perform operations. In some embodiments, the instructions 510C can be executed by the one or more processor(s) 510A to cause the one or more processor(s) 510A to perform operations, such as any of the operations and functions for which the computing system <NUM> and/or the computing device(s) <NUM> are configured, the operations for operating one or more propulsor assemblies (e.g., method <NUM>), as described herein, and/or any other operations or functions of the one or more computing device(s) <NUM>. Accordingly, the method <NUM> may be computerimplemented methods. The instructions 510C can be software written in any suitable programming language or can be implemented in hardware. Additionally, and/or alternatively, the instructions 510C can be executed in logically and/or virtually separate threads on processor(s) 510A. The memory device(s) 510B can further store data 510D that can be accessed by the processor(s) 510A. For example, the data 510D can include data indicative of power flows, data indicative of power demands of various loads in a hybrid propulsion system, data indicative of operational parameters of the hybrid propulsion system, including of a propulsor assemblies of the hybrid propulsion system.

The computing device(s) <NUM> can also include a network interface 510E used to communicate, for example, with the other components of system <NUM> (e.g., via a network). The network interface 510E can include any suitable components for interfacing with one or more network(s), including for example, transmitters, receivers, ports, controllers, antennas, and/or other suitable components. One or more external display devices (not depicted) can be configured to receive one or more commands from the computing device(s) <NUM>.

Although specific features of various embodiments may be shown in some drawings and not in others, this is for convenience only. In accordance with the principles of the present disclosure, any feature of a drawing may be referenced and/or claimed in combination with any feature of any other drawing.

Claim 1:
A hybrid propulsion system (<NUM>) for an aircraft (<NUM>), comprising:
a propulsor assembly (<NUM>) having at least one propulsor (<NUM>);
a power generation assembly comprising at least a first power assembly (<NUM>), a second power assembly (<NUM>), a first electric machine (<NUM>), and a second electric machine (<NUM>), the first power assembly (<NUM>) drivingly coupled to the first electric machine (<NUM>) to produce a first amount of electric power and the second power assembly (<NUM>) drivingly coupled to the second electric machine (<NUM>) to produce a second amount of electric power, wherein the second power assembly (<NUM>) is configured to generate electric power more efficiently than the first power assembly (<NUM>);
a power bus (<NUM>) coupled to the first electric machine (<NUM>) and the second electric machine (<NUM>); and
a controller (<NUM>) operably coupled to the first power assembly (<NUM>), the first electric machine (<NUM>), or both, and to the second power assembly (<NUM>), the second electric machine (<NUM>), or both, the controller (<NUM>) and the power bus (<NUM>) configured to combine at least a portion of the first amount of electrical power and the second amount of electrical power to supply the propulsor assembly (<NUM>),
characterized in that the first power assembly (<NUM>) includes a turbomachine (<NUM>) and the second power assembly (<NUM>) includes an internal combustion engine (<NUM>).