Patent Description:
A conventional commercial aircraft generally includes a fuselage, a pair of wings, and a propulsion system that provides thrust. The propulsion system typically includes at least two aircraft engines, such as turbofan jet engines. Each turbofan jet engine is typically mounted to a respective one of the wings of the aircraft, such as in a suspended position beneath the wing, separated from the wing and fuselage.

Hybrid electric propulsion systems are being developed to improve an efficiency of the conventional commercial aircraft. Various hybrid electric propulsion systems include an electric machine driven by one of the aircraft engines. The inventors of the present disclosure have discovered various configurations and/or methods to address unmet needs for improvements in the known hybrid electric propulsion systems.

<CIT> relates to an aircraft including a parallel hybrid gas turbine electric propulsion system. <CIT> relates to a hybrid electric taxi system or a full electric taxi system. <CIT> relates to a hybrid gas-electric turbine engine. <CIT> relates to an aircraft engine and associated method for driving a fan with a low pressure shaft during taxi operations. In particular it discloses that, in addition to permitting the aircraft to taxi at a desired speed with reduced fuel consumption and brake wear, the taxiing of the aircraft that is provided by an electric taxi system with a core gas turbine engine shut off also serves to cool down the components of the core gas turbine engine.

In one exemplary aspect 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.

The terms "upstream" and "downstream" refer to the relative direction with respect to a flow in a pathway. For example, with respect to a fluid flow, "upstream" refers to the direction from which the fluid flows, and "downstream" refers to the direction to which the fluid flows. However, the terms "upstream" and "downstream" as used herein may also refer to a flow of electricity.

For example, the approximating language may refer to being within a <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, or <NUM> percent margin in either individual values, range(s) of values and/or endpoints defining range(s) of values.

The present disclosure is generally related to a method of operating a hybrid-electric propulsion system for an aircraft during ground operations of the aircraft. For example, in certain exemplary aspects of the method of the present disclosure, the method may generally be operated with a gas turbine engine having an electric machine that is rotatable therewith. The method may generate power with the electric machine through operation of the gas turbine engine during flight operations of the aircraft, storing at least a portion of such electric power in an energy storage unit (e.g., a battery). During or after landing of the aircraft, the gas turbine engine switches from a combustion operating mode to an electric operating mode. Little or no fuel may be provided to the gas turbine engine during the electric operating mode. Further, during the electric operating mode, the electric machine receives power from the electric energy storage unit to drive a low pressure or high pressure system of the gas turbine engine to provide or assist with providing ground operations of the aircraft (e.g., thrust for taxiing the aircraft, rotating a shaft of the engine (including a high pressure shaft and/or a low pressure shaft) for rotor bow mitigation/ soakback mitigation, etc.). In such a manner, the method may generally extract power during operations with power to spare (e.g., during a descent of the aircraft), or during operations that may produce additional power in an efficient manner (e.g., during a cruise of the aircraft), and utilize such power during ground operations.

Referring now to <FIG>, a cross-sectional view of an exemplary embodiment of a gas turbine engine as may incorporate one or more inventive aspects of the present disclosure is provided. In particular, the exemplary gas turbine engine of <FIG> is a configured as a single unducted rotor engine <NUM> defining an axial direction A, a radial direction R, and a circumferential direction C. As is seen from <FIG>, the engine <NUM> takes the form of an open rotor propulsion system and has a rotor assembly <NUM> which includes an array of airfoils arranged around a central longitudinal axis <NUM> of engine <NUM>, and more particularly includes an array of rotor blades <NUM> arranged around the central longitudinal axis <NUM> of engine <NUM>.

Moreover, as will be explained in more detail below, the engine <NUM> additionally includes a non-rotating vane assembly <NUM> positioned aft of the rotor assembly <NUM> (i.e., non-rotating with respect to the central axis <NUM>), which includes an array of airfoils also disposed around central axis <NUM>, and more particularly includes an array of vanes <NUM> disposed around central axis <NUM>.

The rotor blades <NUM> are arranged in typically equally spaced relation around the centerline <NUM>, and each blade has a root <NUM> and a tip <NUM> and a span defined therebetween. Similarly, the vanes <NUM> are also arranged in typically equally spaced relation around the centerline <NUM>, and each has a root <NUM> and a tip <NUM> and a span defined therebetween. The rotor assembly <NUM> further includes a hub <NUM> located forward of the plurality of rotor blades <NUM>.

Additionally, the engine <NUM> includes a turbomachine <NUM> having a core (or high pressure/ high speed system) <NUM> and a low pressure/ low speed system. It will be appreciated that as used herein, the terms "speed" and "pressure" are used with respect to the high pressure/high speed system and low pressure/low speed system interchangeably. Further, it will be appreciate that the terms "high" and "low" are used in this same context to distinguish the two systems, and are not meant to imply any absolute speed and/or pressure values.

The core <NUM> generally includes a high-speed compressor <NUM>, a high speed turbine <NUM>, and a high speed shaft <NUM> extending therebetween and connecting the high speed compressor <NUM> and high speed turbine <NUM>. The high speed compressor <NUM>, the high speed turbine <NUM>, and the high speed shaft <NUM> may collectively be referred to as a high speed spool of the engine. Further, a combustion section <NUM> is located between the high speed compressor <NUM> and high speed turbine <NUM>. The combustion section <NUM> may include one or more configurations for receiving a mixture of fuel and air, and providing a flow of combustion gasses through the high speed turbine <NUM> for driving the high speed spool.

The low speed system similarly includes a low speed turbine <NUM>, a low speed compressor or booster, <NUM>, and a low speed shaft <NUM> extending between and connecting the low speed compressor <NUM> and low speed turbine <NUM>. The low speed compressor <NUM>, the low speed turbine <NUM>, and the low speed shaft <NUM> may collectively be referred to as a low speed spool <NUM> of the engine.

Although the engine <NUM> is depicted with the low speed compressor <NUM> positioned forward of the high speed compressor <NUM>, in certain embodiments the compressors <NUM>, <NUM> may be in an interdigitated arrangement. Additionally, or alternatively, although the engine <NUM> is depicted with the high speed turbine <NUM> positioned forward of the low speed turbine <NUM>, in certain embodiments the turbines <NUM>, <NUM> may similarly be in an interdigitated arrangement.

Referring still to <FIG>, the turbomachine <NUM> is generally encased in a cowl <NUM>. Moreover, it will be appreciated that the cowl <NUM> defines at least in part an inlet <NUM> and an exhaust <NUM>, and includes a turbomachinery flowpath <NUM> extending between the inlet <NUM> and the exhaust <NUM>. The inlet <NUM> is for the embodiment shown an annular or axisymmetric <NUM> degree inlet <NUM> located between the rotor blade assembly <NUM> and the fixed or stationary vane assembly <NUM>, and provides a path for incoming atmospheric air to enter the turbomachinery flowpath <NUM> (and compressors <NUM>, <NUM>, combustion section <NUM>, and turbines <NUM>, <NUM>) inwardly of the guide vanes <NUM> along the radial direction R. Such a location may be advantageous for a variety of reasons, including management of icing performance as well as protecting the inlet <NUM> from various objects and materials as may be encountered in operation.

However, in other embodiments, the inlet <NUM> may be positioned at any other suitable location, e.g., aft of the vane assembly <NUM>, arranged in a non-axisymmetric manner, etc..

As is depicted, the rotor assembly <NUM> is driven by the turbomachine <NUM>, and more specifically, is driven by the low speed spool <NUM>. More specifically, still, engine <NUM> in the embodiment shown in <FIG> includes a power gearbox <NUM>, and the rotor assembly <NUM> is driven by the low speed spool <NUM> of the turbomachine <NUM> across the power gearbox <NUM>. In such a manner, the rotating rotor blades <NUM> of the rotor assembly <NUM> may rotate around the axis <NUM> and generate thrust to propel engine <NUM>, and hence an aircraft to which it is associated, in a forward direction F. For example, in certain embodiments, one or more engines configured in a manner similar to the exemplary engine <NUM> depicted in <FIG> may be incorporated in and utilized with the aircraft of <FIG>, <FIG>, and/or <NUM>.

The power gearbox <NUM> may include a gearset for decreasing a rotational speed of the low speed spool <NUM> relative to the low speed turbine <NUM>, such that the rotor assembly <NUM> may rotate at a slower rotational speed than the low speed spool <NUM>.

As briefly mentioned above the engine <NUM> includes a vane assembly <NUM>. The vane assembly <NUM> extends from the cowl <NUM> and is positioned aft of the rotor assembly <NUM>. The vanes <NUM> of the vane assembly <NUM> may be mounted to a stationary frame or other mounting structure and do not rotate relative to the central axis <NUM>. For reference purposes, <FIG> also depicts the forward direction with arrow F, which in turn defines the forward and aft portions of the system. As shown in <FIG>, the rotor assembly <NUM> is located forward of the turbomachine <NUM> in a "puller" configuration, and the exhaust <NUM> is located aft of the guide vanes <NUM>. As will be appreciated, the vanes <NUM> of the vane assembly <NUM> may be configured for straightening out an airflow (e.g., reducing a swirl in the airflow) from the rotor assembly <NUM> to increase an efficiency of the engine <NUM>. For example, the vanes <NUM> may be sized, shaped, and configured to impart a counteracting swirl to the airflow from the rotor blades <NUM> so that in a downstream direction aft of both rows of airfoils (e.g., blades <NUM>, vanes <NUM>) the airflow has a greatly reduced degree of swirl, which may translate to an increased level of induced efficiency.

Referring still to <FIG>, it may be desirable that the rotor blades <NUM>, the vanes <NUM>, or both, incorporate a pitch change mechanism such that the airfoils (e.g., blades <NUM>, vanes <NUM>, etc.) can be rotated with respect to an axis of pitch rotation either independently or in conjunction with one another. Such pitch change can be utilized to vary thrust and/or swirl effects under various operating conditions, including to adjust a magnitude or direction of thrust produced at the rotor blades <NUM>, or to provide a thrust reversing feature which may be useful in certain operating conditions such as upon landing an aircraft, or to desirably adjust acoustic noise produced at least in part by the rotor blades <NUM>, the vanes <NUM>, or aerodynamic interactions from the rotor blades 16relative to the vanes <NUM>. More specifically, for the embodiment of <FIG>, the rotor assembly <NUM> is depicted with a pitch change mechanism <NUM> for rotating the rotor blades <NUM> about their respective pitch axes <NUM>, and the vane assembly <NUM> is depicted with a pitch change mechanism <NUM> for rotating the vanes <NUM> about their respective pitch axes <NUM>.

It will be appreciated, however, that the exemplary single rotor unducted engine <NUM> depicted in <FIG> is by way of example only, and that in other exemplary embodiments, the engine <NUM> may have any other suitable configuration, including, for example, any other suitable number of shafts or spools, turbines, compressors, etc.; fixed-pitch blades <NUM>, <NUM>, or both; a direct-drive configuration (i.e., may not include the gearbox <NUM>); etc..

Additionally, or alternatively, in other exemplary embodiments, any other suitable gas turbine engine may be provided. For example, in other exemplary embodiments, the gas turbine engine may be a ducted turbofan engine, a turboshaft engine, a turboprop engine, turbojet engine, etc. Moreover, for example, although the engine is depicted as a single unducted rotor engine, in other embodiments, the engine may include a multi-stage open rotor configuration, and aspects of the disclosure described hereinbelow may be incorporated therein.

Further, still, in other exemplary embodiments, the engine <NUM> may be configured as a ducted turbofan engine. For example, referring briefly to <FIG>, an engine <NUM> in accordance with another exemplary embodiment of the present disclosure is depicted. The exemplary embodiment of <FIG> may be configured in substantially the same manner as the exemplary engine <NUM> described above with respect to <FIG>, and the same or similar reference numerals may refer to the same or similar parts. However, as will be appreciated, for the embodiment shown, the engine <NUM> further includes a nacelle <NUM> circumferentially surrounding at least in part the rotor assembly <NUM> and turbomachine <NUM>, defining a bypass passage <NUM> therebetween.

Referring now back to <FIG>, it will further be appreciated that the engine is integrated with an electric power system <NUM>. The electric power system <NUM> generally includes an electric machine <NUM> coupled to at least the low pressure system, an energy storage unit <NUM>, and for the embodiment depicted, an auxiliary power unit <NUM>. The auxiliary power unit <NUM> may include a combustion engine driving an electric generator, and may be located remotely from the engine <NUM>. For example, in at least certain exemplary embodiments, the auxiliary power unit <NUM> may be located within a fuselage of the aircraft utilizing the engine <NUM>, e.g., at an aft end of the aircraft (see, e.g., <FIG>).

Further, for the embodiment shown, the electric power system <NUM> includes an electric power bus <NUM> electrically connecting the various components of electric power system <NUM>. The electric power bus <NUM> may be, e.g., one or more electrical lines arranged in any suitable configuration.

Further, still, for the embodiment shown, the electric machine <NUM> of the electric power system <NUM> is a low speed ("LS") electric machine 102A coupled to the low pressure system of the engine. More specifically, for the embodiment shown, the LS electric machine 102A is embedded within the engine <NUM>, at a location within or aft of the turbine section of the engine <NUM>, and inward of the core airflow path <NUM> through the engine <NUM> along the radial direction R. It will be appreciated, however, that in other example embodiments, the LS electric machine 102A may additionally, or alternatively, be configured in the other suitable manner. For example, in other embodiments, the LS electric machine 102A may be embedded within a compressor section of the engine <NUM>, may be located outward of core airflow path <NUM> along the radial direction R (and, e.g., within the cowl <NUM>), may be driven off of a gearbox (such as an accessory gearbox), etc..

Moreover, for the embodiment shown, the LS electric machine 102A is not the only electric machine <NUM> of the electric power system <NUM> integrated with the engine <NUM>. More specifically, the electric power system <NUM> further includes a high speed ("HS") electric machine 102B coupled to the high-pressure system/core of the engine <NUM>, and in electrical communication with the electric power bus <NUM>. The HS electric machine 102B is, for the embodiment shown, also embedded within the engine <NUM> at a location inward of the core airflow path <NUM>. However, for the embodiment shown, the HS electric machine 102B is located within the compressor section of the engine <NUM>. It will be appreciated that in other embodiments, the HS electric machine 102B may alternatively be positioned outward of the core airflow path <NUM> along the radial direction R, driven through, e.g., a geared connection. For example, in certain embodiments, the HS electric machine 102B may be coupled to an accessory gearbox (not shown), which is in turn coupled to the high-pressure system of the engine <NUM>.

In at least certain exemplary embodiments, the energy storage unit <NUM> may include one or more batteries. Additionally, or alternatively, the energy storage unit <NUM> may include one or more supercapacitor arrays, one or more ultracapacitor arrays, or both. In at least certain embodiments, the energy storage unit <NUM> may be configured to provide at least <NUM> kilowatts (kW) of energy to the electric power system <NUM>, such as at least <NUM> kW, such as at least <NUM> kW, such as at least <NUM> kW, such as at least <NUM> kW, such as at least <NUM> kW, such as at least <NUM> kW, such as at least <NUM> kW, such as up to <NUM> megawatts (MW), such as up to <NUM> megawatts (MW). In one or more of these configurations, the amount of power provided may refer to a peak power output at any instant time during a discharge period. Further, the energy storage unit <NUM> may be configured to provide such electrical power for at least two minutes, such as at least three minutes, such as at least five minutes, such as up to an hour. For example, the energy storage unit <NUM> may be configured to store at least <NUM> kW-minutes of power, such as at least <NUM> kW-minutes of power, such as at least <NUM> kW-minutes of power, such as at least <NUM> kW-minutes of power, such as at least <NUM> kW-minutes of power, such as at least <NUM> kW-minutes of power, such as up to <NUM> MW-minutes of power. In one or more of these configurations, the amount of power provided above may be a peak power output at any instant time during a discharge period.

As is also depicted in <FIG>, the engine <NUM> includes one or more accessory systems. Specifically, for the embodiment shown, the engine <NUM> includes a blower <NUM> and a lubrication system <NUM>. The blower <NUM> is in the embodiment shown positioned within an undercowl area (an area under the cowl <NUM> and outward of the turbomachinery flowpath <NUM>) and is electrically connected to the energy storage unit. The blower <NUM> may be configured to provide an airflow to the compressor section, the combustion section, or both from, e.g., an ambient location (such as ambient air drawn through a bypass duct), air drawn from an interior space of the aircraft wing, a location under the cowl <NUM>, or a combination of these locations. The blower <NUM> may be operated after the engine <NUM> is shut down to maintain an airflow through certain components of the engine <NUM> to prevent or minimize heat within various components of the engine <NUM> from "soaking back" into, e.g., fuel nozzles of the combustion section and heating such fuel nozzles above a temperature which would cause coking of any remaining fuel therein. For example, in certain exemplary aspects, the blower may produce an airflow through the turbomachinery flowpath <NUM>, the undercowl area, or both. In particular, the blower <NUM> may provide an airflow from the undercowl area through a compressor or combustor port and into the turbomachinery flowpath <NUM> through the compressor and combustion sections of the engine <NUM> (as is indicated in phantom lines in <FIG>). Alternatively, the blower <NUM> may pull an airflow from the turbomachinery flowpath <NUM> through the compressor and/or combustor sections of the engine <NUM> to provide a desired soak-back protection.

Further, as noted, for the embodiment shown, the engine <NUM> includes a lubrication system <NUM>. The lubrication system <NUM> may be a lubrication system for the low pressure system, such that when driven by the power provided from the electric power system <NUM>, the lubrication system <NUM> circulates a lubrication fluid through various portions of the low pressure system (e.g., bearings, sumps, heat exchangers, etc.). In such a manner, the lubrication system <NUM> may continue to support the lubrication functionality of the engine <NUM> for the low pressure system when, e.g., the LS electric machine 102A rotates the low pressure system when the engine <NUM> is in an electric operating mode, and a fuel flow to the combustion section has been stopped.

It will be appreciated, however, in another example embodiment, the lubrication system <NUM> may additionally or alternatively include a lubrication system for the high-pressure system, such that when driven by the power provided by the electric power system <NUM> the lubrication system <NUM> circulates a lubrication fluid through various portions of the high-pressure system (e.g., bearings, sumps, heat exchangers, etc.).

Referring still to <FIG>, the exemplary electric power system <NUM> is operably connected to a controller <NUM>. The controller <NUM> may be an engine controller for the engine <NUM> (e.g., a Full Authority Digital Engine Control controller), may be an aircraft controller, may be a controller dedicated to the electric power system <NUM>, etc..

The controller <NUM> may be configured to receive data indicative of various operating conditions and parameters of the engine <NUM> during operation of the engine <NUM>. For example, as will be appreciated from <FIG>, the engine <NUM> includes one or more sensors <NUM> configured to sense data indicative of various operating conditions and parameters of the engine <NUM>, such as rotational speeds, temperatures, pressures, vibrations, etc. For example, the one or more sensors <NUM> may sense data indicative of a temperature parameter within the engine <NUM>, such as an exhaust gas temperature, a combustion section temperature, a compressor exit temperature, etc. Additionally, or alternatively, the one or more sensors <NUM> may sense data indicative of a speed of the engine <NUM>, such as a rotational speed of the low pressure system, a rotational speed of the high-pressure system, a rotational speed of the rotor section <NUM>, etc. It will be appreciated that although a single sensor <NUM> is depicted in <FIG>, a plurality of sensors <NUM> providing the above functionality may be positioned throughout the engine <NUM> to sense the relevant data. In addition, as will be appreciated from the description herein, the controller <NUM> may also be configured to receive data form other sources, such as from an aircraft incorporating the engine, such as from one or more sensors of the aircraft incorporating the engine. In such a manner, the controller <NUM> may receive data indicative of an altitude of the aircraft, a signal to engage a supplemental power from e.g., a pilot or other operator, etc..

Referring particularly to the operation of the controller <NUM>, in at least certain embodiments, the controller <NUM> can include one or more computing device(s) <NUM>. The computing device(s) <NUM> can include one or more processor(s) 118A and one or more memory device(s) 118B. The one or more processor(s) 118A 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) 118B 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) 118B can store information accessible by the one or more processor(s) 118A, including computer-readable instructions 118C that can be executed by the one or more processor(s) 118A. The instructions 118C can be any set of instructions that when executed by the one or more processor(s) 118A, cause the one or more processor(s) 118A to perform operations. In some embodiments, the instructions 118C can be executed by the one or more processor(s) 118A to cause the one or more processor(s) 118A to perform operations, such as any of the operations and functions for which the controller <NUM> and/or the computing device(s) <NUM> are configured, the operations for operating an electric power system <NUM> (e.g., method <NUM>), as described herein, and/or any other operations or functions of the one or more computing device(s) <NUM>. The instructions 118C can be software written in any suitable programming language or can be implemented in hardware. Additionally, and/or alternatively, the instructions 118C can be executed in logically and/or virtually separate threads on processor(s) 118A. The memory device(s) 118B can further store data 118D that can be accessed by the processor(s) 118A. For example, the data 118D can include data indicative of power flows, data indicative of engine <NUM>/ aircraft operating conditions, and/or any other data and/or information described herein.

The computing device(s) <NUM> can also include a network interface 118E used to communicate, for example, with the other components of the engine <NUM>, the aircraft incorporating the engine <NUM>, the electric power system <NUM>, etc. For example, in the embodiment depicted, as noted above, the engine <NUM> includes one or more sensors <NUM> for sensing data indicative of one or more parameters of the engine <NUM> and various accessory systems, and the electric power system <NUM> includes an energy storage unit <NUM>, an LS electric machine 102A, an HS electric machine 102B, and an auxiliary power unit <NUM>. The controller <NUM> is operably coupled to these components through, e.g., the network interface 118E, such that the controller <NUM> may receive data indicative of various operating parameters sensed by the one or more sensors <NUM> during operation, various operating conditions of the components, etc., and further may provide commands to control electrical flow of the electric power system <NUM> and other operating parameters of these systems, e.g., in response to the data sensed by the one or more sensors <NUM> and other conditions.

The network interface 118E 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. For example, in the embodiment shown, the network interface 118E is configured as a wireless communication network wirelessly in communication with these components (as is indicated by the dashed communication lines in <FIG>).

Referring now to <FIG>, a flow diagram of a method <NUM> for operating a hybrid electric propulsion system of an aircraft is provided. The method <NUM> may be utilized with one or more of the exemplary embodiments described above with reference to <FIG>, and/or with one or more of the other exemplary embodiments described herein. The hybrid electric propulsion system includes a gas turbine engine having a high-pressure system, a low pressure system, an electric machine coupled to at least the low-pressure system, and an energy storage unit.

The method <NUM> includes at (<NUM>) operating the electric machine as an electric generator to charge the energy storage unit during a flight operation of the aircraft. For example, operating the electric machine as an electric generator to charge the energy storage unit at (<NUM>) may include operating the electric machine as an electric generator to charge the energy storage unit during a cruise operation of the aircraft, a descent operation of the aircraft, a climb operation of the aircraft, etc..

The method <NUM> further includes at (<NUM>) switching the gas turbine engine to an electric operating mode during or after a landing operation of the aircraft. In certain exemplary aspects, switching the gas turbine engine to an electric operating mode at (<NUM>) may include at (<NUM>) shutting off a fuel flow to the gas turbine engine. Moreover, in certain exemplary aspects, switching the gas turbine engine to an electric operating mode at (<NUM>) may include at (<NUM>) switching the gas turbine engine to an electric operating mode after a landing operation of the aircraft. The landing operation of the aircraft may be considered complete once all wheels of the aircraft are on the ground following a descent of the aircraft. Additionally, or alternatively, the landing operation of the aircraft may be considered complete once a vehicle ground speed falls below about <NUM> kilometres per hour (about <NUM> miles per hour), once a command is received from the pilot or operator that the landing operation is complete, in response to a position of the vehicle (e.g., once the vehicle is off the runway), etc..

The method <NUM> further includes driving a system of the gas turbine engine using power from the energy storage unit while in the electric operating mode to provide or assist with providing ground operations the aircraft. More specifically, the method <NUM> further includes at (<NUM>) driving the low pressure system, the high pressure system, or both with the electric machine using power from the energy storage unit while in the electric operating mode to provide or assist with providing ground operations the aircraft. The term "ground operations" generally refer to any operations from the time the aircraft touches down (i.e., wheels touch the ground) to when all engine operation is stopped (including any accessory systems that operate after the engine stops rotating and receiving fuel). As will be appreciated from the discussion herein, the ground operations include driving the low pressure system of the aircraft to mitigate engine soakback, and may include taxiing the aircraft, rotating one or more aspects of the engine to prevent or mitigate rotor bow, driving accessory systems of the engine, or the like.

More specifically, for the exemplary aspect depicted, driving the low pressure system, the high pressure system, or both with the electric machine using power from the energy storage unit while in the electric operating mode to provide or assist with providing ground operations the aircraft at (<NUM>) includes at (<NUM>) driving the low pressure system, the high-pressure system, or both with the electric machine using power from the energy storage unit while in the electric operating mode to generate thrust for the aircraft with the gas turbine engine for taxiing the aircraft. More specifically, for the exemplary aspect depicted, driving the low-pressure system, the high-pressure system, or both with the electric machine using power from the energy storage unit to generate thrust for the aircraft for taxiing the aircraft at (<NUM>) includes at (<NUM>) driving the low-pressure system with the electric machine. Notably, driving the low-pressure system with the electric machine at (<NUM>) may correspondingly rotate a rotor assembly of the engine (such as, e.g., a fan assembly, a propeller, or the like, such as the rotor assembly <NUM> of the exemplary engine <NUM> of <FIG>) to generate thrust for the aircraft for taxiing the aircraft.

In accordance with the above, it will be appreciated that in at least certain exemplary embodiments, in order to provide a desired amount of thrust for taxiing the aircraft, driving the low-pressure system with the electric machine at (<NUM>) may include at (<NUM>) providing at least about <NUM> hp to the gas turbine engine with the electric machine to generate thrust for the aircraft. For example, in certain exemplary aspects, driving the low-pressure system with the electric machine at (<NUM>) may include providing at least about <NUM> hp to the gas turbine engine with the electric machine, such as at least about <NUM> hp, such as up to about <NUM> hp. Moreover, in at least certain exemplary aspects, driving the low-pressure system with the electric machine at (<NUM>) may include at (<NUM>) providing at least about <NUM> hp for the gas turbine engine with the electric machine to generate thrust for the aircraft for at least about three minutes, such as for at least about five minutes, such as for at least about <NUM> minutes, such as for at least about <NUM> minutes.

Referring still to the exemplary aspect of the method <NUM> depicted in <FIG>, it will be appreciated that the exemplary method <NUM> may further be utilized to reduce a negative effect of heat soak back within the gas turbine engine following a shutdown of the gas turbine engine. The exemplary method <NUM> depicted further includes at (<NUM>) parking the aircraft. Parking the aircraft at (<NUM>) generally refers to placing the aircraft in a condition without movement for a prolonged period of time. As will be appreciated, once the aircraft is parked, an airflow through the gas turbine engines of the aircraft may be reduced, potentially creating an opportunity for residual heat within the engine to transfer to certain components, potentially damaging such components. For example, residual heat may transfer from the various rotors and core components of the high pressure system and combustion section to fuel nozzles of the combustion section, to accessories such as control units and other electronics, etc. Additionally or alternatively, the residual heat may transfer to fuel at locations other than the fuel nozzles, potentially heating the fuel near or inside the gas turbine engine and resulting in coke formation.

Accordingly, for the exemplary aspect of the method <NUM> depicted in <FIG>, driving the low pressure system, the high pressure system, or both with the electric machine using power from the energy storage unit while in the electric operating mode to provide or assist with providing ground operations the aircraft at (<NUM>) additionally includes at (<NUM>) driving the low-pressure system of the aircraft with the electric machine to mitigate engine soak-back. Notably, driving the low-pressure system of the aircraft with the electric machine to mitigate engine soak-back at (<NUM>) includes at (<NUM>) driving the low-pressure system of the aircraft with the electric machine to mitigate engine soak-back after parking the aircraft at (<NUM>).

In such a manner, driving the low-pressure system of the aircraft with the electric machine at (<NUM>) may include inducing and airflow through a core air flow path of the engine to reduce engine soak-back by providing the airflow through the core. Additionally, or alternatively, driving the low-pressure system of the aircraft with the electric machine at (<NUM>) may include inducing and airflow through a core air flow path of the engine to reduce engine soak-back by providing an airflow that rotates the core.

For example, driving the low-pressure system of the aircraft with the electric machine to mitigate engine soak back at (<NUM>) further includes at (<NUM>) rotating the low-pressure system with the electric machine at a rotational speed of less than <NUM> revolutions per minute and greater than one revolution per minute. For example, in certain exemplary aspects, driving the low-pressure system of the engine with the electric machine to mitigate engine soak back at (<NUM>) may include rotating the low-pressure system of the engine with the electric machine at a rotational speed of less than <NUM> revolutions per minute, such as at a rotational speed of less than <NUM> revolutions per minute, such as a rotational speed less than <NUM> revolutions per minute. Operating the engine in accordance with one or more these exemplary aspects may provide the desired amount of soak back mitigation for the engine.

Further, it will be appreciated that for the embodiment depicted, driving the low-pressure system of the aircraft with the electric machine to mitigate engine soak back at (<NUM>) may include at (<NUM>) driving the low-pressure system of the aircraft with the electric machine to mitigate engine soak back in response to a sensed condition, or in response to any other operation condition. For example, the sensed condition may be a temperature parameter of the engine (e.g., data indicative of an exhaust gas temperature, a compressor exit temperature, etc.), a time parameter (e.g., a time since shutdown of the engine or parking of the aircraft, etc.), or any other suitable sensed condition. The sensed condition may be a condition sensed with a controller from one or more engine sensors. The operation condition may include the sensed condition(s) as well as, e.g., a manual signal received from an operator.

Moreover, it will be appreciated that the method <NUM> may further include rotating a high pressure system of the engine with an electric machine at a relatively slow speed to reduce a bowed rotor condition. For example, the method <NUM> may rotate a high pressure system at a rotational speed less than about <NUM> revolutions per minute, such as a rotational speed less than <NUM> revolutions per minute, such as a rotational speed less than <NUM> revolutions per minute, such as at rotational speed less than <NUM> revolution per minute. Such a process step may allow for the heat soak to be evenly distributed around a circumference of, e.g., one or more rotors of the high pressure system, potentially reducing or eliminating a rotor bowed condition.

It will be appreciated, however, that such efforts may, depending on, e.g., certain flight operating conditions and engine operating conditions during the preceding flight, and/or certain environmental conditions, may not be sufficient to mitigate engine soak back to the extent desired. For the exemplary aspect depicted, driving the low-pressure system, the high-pressure system, or both at (<NUM>) further includes at (<NUM>) receiving data indicative of an engine temperature parameter being in excess of a predetermined threshold while driving the low-pressure system of the aircraft with the electric machine to mitigate engine soak-back at (<NUM>). Receiving data at (<NUM>) may include receiving data from one or more engine sensors with a controller.

In response, driving the low-pressure system, the high-pressure system, or both at (<NUM>) additionally includes at (<NUM>) driving the high-pressure system of the gas turbine engine with a second electric machine to increase a cooling of the gas turbine engine (in response to receiving the data at (<NUM>)). In at least certain exemplary aspects, driving the high-pressure system of the gas turbine engine with a second electric machine at (<NUM>) may include rotating the high-pressure system of the gas turbine engine with the second electric machine at a rotational speed of less than <NUM> revolutions per minute and greater than one revolution per minute.

Notably, however, in other exemplary aspects, instead of receiving the engine temperature parameter at (<NUM>), the method <NUM> may receive data indicative of an operation condition while driving the low pressure system of the aircraft with the electric machine to mitigate engine soak-back, and the method <NUM> may further include driving the high pressure system of the gas turbine engine with a second electric machine to increase a cooling of the gas turbine engine in response to receiving data indicative of the operation condition. The operation condition may be a time parameter, a temperature parameter, a manual signal, or a combination thereof.

Further, it will be appreciated that for the exemplary aspect depicted, driving the low-pressure system of the gas turbine engine with the electric machine to mitigate engine soak back at (<NUM>) may include driving the low-pressure system of electric machine for a time period following parking the aircraft at (<NUM>) between at least about two minutes and two hours, such as equal to at least five minutes, such as equal to at least <NUM> minutes, such as up to <NUM> minutes, such as up to one hour.

Referring now to <FIG>, additional example aspects of the method <NUM> are depicted. The method <NUM> depicted in <FIG> may be configured in substantially the same manner as the method <NUM> depicted in <FIG>. However, for the exemplary aspect of <FIG>, the method <NUM> further includes at (<NUM>) providing power from the energy storage unit to an accessory system the gas turbine engine after switching the gas turbine engine to the electric operating mode. In at least certain exemplary aspects, providing power from the energy storage unit to an accessory system of the gas turbine engine at (<NUM>) may include providing power to an accessory engine cooling system (such as a core soak back blower), to a lubrication system (such as a lubrication system for the low-pressure system, a lubrication system for the high-pressure system, or both), etc. In these examples the accessory systems may alternatively, selectively, or additionally be configured to operate during or after shutdown of the engine as in, e.g., once the aircraft has arrived at the gate and/or parked. For example, an electric blower located within an undercowl area and operated after or during engine shutdown, may activate and operate for a predetermined amount of time based upon relevant operating or environmental parameters (e.g., taxi time, ambient air temperature) to limit a peak soakback temperature (e.g., <NUM> Deg. F to avoid fuel coking, or <NUM> Deg. C to avoid damage to an electrical machine located near the LP turbine, if provided), or the blower may operate for a period of time based on a sensed temperature located at/near, e.g., combustor nozzle, T3 location, or aft end of LP turbine. The blower may be configured to produce a forced air stream through the core (e.g., by way of a bleed port, plenum, etc. located at the downstream end of compressor) by drawing air from either a third stream of the engine, booster bypass valve or through the undercowl area (drawing air through the aft end of engine).

Additionally, although not depicted, it will be appreciated that in still certain exemplary aspects, the method <NUM> may further include providing power from the energy storage unit to one or more accessory systems of the aircraft, such as to one or more electronic control systems, environmental control systems, variable geometry control systems, hydraulic systems, pneumatic systems, etc..

Further, in one or more of these exemplary aspects, the energy storage unit may not be sufficient to meet all of the electrical power needs for the gas turbine engine, the electric machine, and/or aircraft. With such exemplary aspect, the method <NUM> further includes at (<NUM>) determining a power need for the electric machine, the gas turbine engine, the aircraft, or a combination thereof after switching the gas turbine engine to the electric operating mode is in excess of an available power of the energy storage unit, and at (<NUM>) operating an auxiliary power unit of the aircraft to generate an additional amount of power for the electric machine, the gas turbine engine, the aircraft, or a combination thereof.

In at least certain exemplary aspects, determining the power need for the electric machine, the gas turbine engine, the aircraft, or a combination thereof at (<NUM>) may include determining a power need of the electric machine after switching the gas turbine engine to the electric operating mode is in excess of the available power of the energy storage unit for the electric machine. Further, determining the power need may include, e.g., estimating the time remaining for the electric machine to drive the low-pressure system of the gas turbine engine to, e.g., mitigate engine soak pack, taxi the aircraft, etc.; receiving temperature data of the electric machine to estimate an additional amount of time to rotate the low-pressure system of the gas turbine engine, the high-pressure system of the gas turbine engine, or both to mitigate engine soak back; receiving temperature data of the electric machine indicative of a need to operate additional aircraft cooling systems (such as, e.g., the engine soak back blower, a second electric machine to rotate the high-pressure system, etc.); or the like.

Further, operating the auxiliary power unit of the aircraft to generate an additional amount of power for the electric machine, the gas turbine engine, the aircraft, or a combination thereof at (<NUM>) may include providing electrical power from the auxiliary power unit to the energy storage unit and/or the electric machine via an electric power bus of the power system.

Further, still, it will be appreciated that as noted above, the ground operations described herein are powered by the electric machine driving the low pressure system, the high pressure system, or both with power from the energy storage unit while in the electric operating mode. In at least certain exemplary aspects a substantial portion of all of the electric power utilized by the electric machine while providing or assisting with providing the ground operations ("Total Electrical Power") may come from the electric energy storage unit. For example, at least <NUM>% of the Total Electrical Power may come from the electric energy storage unit, such as at least about <NUM>%, such as at least about <NUM>%, such as at least about <NUM>%, such as at least about <NUM>%, such as at least about <NUM>% of the Total Electrical Power may come from the electric energy storage unit.

However, in certain exemplary aspects not all of the Total Electrical Power may come from the electric energy storage unit. For example, in certain exemplary aspects, up to <NUM>% of the Total Electrical Power may come from the electric energy storage unit, such as up to <NUM>% of the Total Electrical Power may come from the electric energy storage unit. The remainder of the Total Electrical Power may be provided from the APU, and/or depending on the ground operation, from a power source external to the aircraft (e.g., a ground power source).

For example, in certain exemplary aspects, when the ground operations include taxiing the aircraft, the energy storage unit may provide all the power for taxiing the aircraft, and then supplemental power may be provided to rotate the engine to mitigate rotor bow/ soakback, or vice versa. Alternatively, supplemental power may be provided for a peak output power need (such as during taxiing), but not for a lower power output need (such as rotating the engine to mitigate rotor bow/ soakback).

Claim 1:
A method (<NUM>) for operating a hybrid-electric propulsion system of an aircraft, the hybrid-electric propulsion system comprising a gas turbine engine having a high pressure system, a low pressure system, an electric machine coupled to at least the low pressure system, and an energy storage unit, the method comprising:
operating (<NUM>) the electric machine as an electric generator to charge the energy storage unit during a flight operation of the aircraft;
switching (<NUM>) the gas turbine engine to an electric operating mode during or after a landing operation of the aircraft;
driving (<NUM>) a system of the gas turbine engine with the electric machine using power from the energy storage unit while in the electric operating mode to provide or assist with providing ground operations the aircraft;
wherein driving the system of the gas turbine engine comprises driving the low pressure system, the high pressure system, or both, wherein the ground operations include driving (<NUM>)
the low pressure system of the aircraft with the electric machine to mitigate engine soak-back, and wherein the method further comprises:
parking (<NUM>) the aircraft, wherein driving the low pressure system of the aircraft with the electric machine to mitigate engine soak-back comprises driving (<NUM>) the low pressure system of the aircraft with the electric machine to mitigate engine soak-back after parking the aircraft.