Ground operation of hybrid propulsion system

Abstract

A method for operating a hybrid propulsion system of an aircraft is provided. The hybrid propulsion system includes a gas turbine engine having a high pressure system and a low pressure system, an electric machine coupled with at least one of the high pressure system or the low pressure system, and an energy storage unit. The method includes operating the electric machine as a generator during flight operation of the aircraft, charging the energy storage unit; switching the gas turbine engine to an electrical mode of operation during or after a landing operation of the aircraft; in an electrical mode of operation, a system for driving a gas turbine engine using electric power from an energy storage unit to provide or assist in providing ground operation of an aircraft.

Description

[0001] This application is a divisional application of the patent application with the application number 202111006435.3 and the invention title "Ground Operation of a Hybrid Propulsion System" filed on August 30, 2021.

[0002] Cross - Reference to Related Applications

[0003] This application is a non - provisional application and claims priority under 35 U.S.C.§119(e) to U.S. Provisional Application No. 63 / 072,573, filed on August 31, 2020, the entire content of which is incorporated herein by reference. Technical Field

[0004] This subject matter generally relates to hybrid aircraft propulsion systems, hybrid systems for aircraft and aircraft engines, and methods for operating them. Background Art

[0005] Conventional commercial aircraft typically include a fuselage, a pair of wings, and a propulsion system that provides thrust. The propulsion system generally includes at least two aircraft engines, such as turbofan jet engines. Each turbofan jet engine is typically mounted on a respective wing of the aircraft, for example, in a suspended position under the wing, separate from the wing and the fuselage.

[0006] Hybrid propulsion systems are being developed to improve the efficiency of conventional commercial aircraft. Various hybrid propulsion systems include an electric motor driven by one of the aircraft engines. The inventors of the present invention have discovered various configurations and / or methods to address the unmet needs for improvements to known hybrid propulsion systems. Summary of the Invention

[0007] Aspects and advantages of the present invention will be set forth in part in the following description, or may be apparent from the description, or may be learned by practice of the present invention.

[0008] In an exemplary aspect of the present disclosure, a method for operating a hybrid propulsion system of an aircraft is provided. The hybrid propulsion system includes a gas turbine engine having a high - pressure system, a low - pressure system, an electric motor coupled to at least one of the high - pressure system or the low - pressure system, and an energy storage unit. The method includes operating the electric motor as a generator during a flight operation of the aircraft to charge the energy storage unit; switching the gas turbine engine to an electric operation mode during or after a landing operation of the aircraft; and using electric power from the energy storage unit to drive the systems of the gas turbine engine in the electric operation mode to provide or assist in providing ground operation of the aircraft.

[0009] These and other features, aspects, and advantages of the present invention will become better understood with reference to the following description and the appended claims. The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the invention and, together with the description, serve to explain the principles of the invention. BRIEF DESCRIPTION OF THE DRAWINGS

[0010] A complete and enabling disclosure of the present invention, including the best mode thereof, to one of ordinary skill in the art, is set forth in the specification, which makes reference to the accompanying drawings, in which:

[0011] Figure 1 is a schematic cross-sectional view of a gas turbine engine in accordance with an exemplary embodiment of the present disclosure.

[0012] Figure 2 is a schematic cross-sectional view of a gas turbine engine in accordance with another exemplary embodiment of the present disclosure.

[0013] Figure 3 is a flowchart of a method for operating a gas turbine engine in accordance with an exemplary aspect of the present disclosure.

[0014] Figure 4 is a flowchart of a method for operating a gas turbine engine in accordance with another exemplary aspect of the present disclosure. DETAILED DESCRIPTION

[0015] Reference will now be made in detail to embodiments of the invention, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerical and letter designations to refer to features in the drawings. In the drawings and the description, like or similar designations have been used to refer to like or similar parts of the invention.

[0016] As used herein, the terms "first," "second," and "third" may be used interchangeably to distinguish one component from another, rather than to denote the position or importance of individual components.

[0017] The terms "front" and "rear" refer to relative positions within a gas turbine engine or vehicle and refer to the normal operating attitude of the gas turbine engine or vehicle. For example, with respect to a gas turbine engine, the front refers to the position near the engine inlet, and the rear refers to the position near the engine nozzle or exhaust.

[0018] The terms "upstream" and "downstream" refer to the relative direction with respect to the flow in a path. For example, with respect to fluid flow, "upstream" refers to the direction from which the fluid flows, and "downstream" refers to the direction in which the fluid flows. However, the terms "upstream" and "downstream" as used herein may also refer to the flow of electricity.

[0019] The singular forms "a", "an", and "the" include plural references unless the context clearly dictates otherwise.

[0020] As used herein throughout the specification and claims, approximate language is applied to modify any quantitative representation that could permissibly vary without resulting in a change in the basic function associated therewith. Thus, a numerical value modified by a term or terms, such as "about", "approximately", and "substantially", is not limited to the precise value specified. At least in some instances, the approximate language may correspond to the precision of the instrument used to measure the value, or the precision of the method or machine of constructing or manufacturing the component and / or system. At least in some instances, the approximate language may correspond to the precision of the instrument for measuring a numerical value, or the precision of the method or machine of constructing or manufacturing the component and / or system. For example, the approximate language may refer to within 1, 2, 4, 5, 10, 15, or 20% of a single numerical value, a numerical range, and / or the endpoints defining the numerical range.

[0021] Herein, and throughout the specification and claims, range limitations are combined and interchanged, such ranges are identified and include all subranges subsumed therein unless the context or language indicates otherwise. For example, all ranges disclosed herein include the endpoints, and the endpoints may be combined independently of each other.

[0022] The present invention generally relates to a method of operating a hybrid propulsion system of an aircraft during ground operation of the aircraft. For example, in certain exemplary aspects of the methods of the present disclosure, the method may generally be operated with a gas turbine engine having an electric machine rotatable therewith. The method may generate electrical power using the electric machine by operating the gas turbine engine during flight operation of the aircraft and store at least a portion of such electrical power in an energy storage unit (e.g., a battery). During or after landing of the aircraft, the gas turbine engine may be switched from a combustion operation mode to an electric operation mode. During the electric operation mode, little or no fuel may be supplied to the gas turbine engine. Additionally, during the electric operation mode, the electric machine may receive electrical power from the electrical energy storage unit to drive a low-pressure or high-pressure system of the gas turbine engine to provide or assist in providing ground operation of the aircraft (e.g., thrust for taxiing the aircraft, rotating the shaft of the engine (including the high-pressure shaft and / or the low-pressure shaft) for rotor bow relief / soak mitigation, etc.). In this manner, the method may generally extract electrical power during operation (e.g., during descent of the aircraft) with backup power, or during an operation where additional electrical power can be generated in an efficient manner (e.g., during cruise of the aircraft), and utilize such electrical power during ground operation.

[0023] Now referring to Figure 1 , there is provided a cross-sectional view of an exemplary embodiment of a gas turbine engine that may incorporate one or more inventive aspects of the present disclosure. In particular,Figure 1 An exemplary gas turbine engine is configured as a single non-ducted rotor engine 10 that defines an axial direction A, a radial direction R, and a circumferential direction C. As seen from Figure 1 it can be seen that the engine 10 takes the form of an open rotor propulsion system and has a rotor assembly 12 that includes an array of airfoils disposed about the central longitudinal axis 14 of the engine 10, and more particularly includes an array of rotor blades 16 disposed about the central longitudinal axis 14 of the engine 10.

[0024] In addition, as will be explained in more detail below, the engine 10 further includes a non-rotating stator assembly 18 (i.e., non-rotating relative to the central axis 14) positioned behind the rotor assembly 12, which includes an array of airfoils also disposed about the central axis 14, and more specifically includes an array of stators 20 disposed about the central axis 14.

[0025] The rotor blades 16 are disposed about the centerline 14 in a typically equally spaced relationship, and each blade has a root 22 and a tip 24 and a span defined therebetween. Similarly, the stators 20 are also disposed about the centerline 14 in a typically equally spaced relationship, and each stator has a root 26 and a tip 28 and a span defined therebetween. The rotor assembly 12 further includes a hub 44 located in front of the plurality of rotor blades 16.

[0026] In addition, the engine 10 includes a turbine 30 having a core (or high-pressure / high-speed system) 32 and a low-pressure / low-speed system. It will be understood that, as used herein, the terms "speed" and "pressure" are used interchangeably with respect to the high-pressure / high-speed system and the low-pressure / low-speed system. In addition, it will be understood that the terms "high" and "low" are used in the same context to distinguish between the two systems and do not imply any absolute speed and / or pressure values.

[0027] The core 32 generally includes a high-speed compressor 34, a high-speed turbine 36, and a high-speed shaft 38 extending therebetween and connecting the high-speed compressor 34 and the high-speed turbine 36. The high-speed compressor 34, the high-speed turbine 36, and the high-speed shaft 38 may be collectively referred to as the high-speed spool of the engine. In addition, a combustion section 40 is located between the high-speed compressor 34 and the high-speed turbine 36. The combustion section 40 may include one or more configurations for receiving a mixture of fuel and air and providing combustion gases to flow through the high-speed turbine 36 to drive the high-speed spool.

[0028] The low-speed system similarly includes a low-speed turbine 42, a low-speed compressor or supercharger 44, and a low-speed shaft 46 extending therebetween and connecting the low-speed compressor 44 and the low-speed turbine 42. The low-speed compressor 44, the low-speed turbine 42, and the low-speed shaft 46 may be collectively referred to as the low-speed spool 55 of the engine.

[0029] Although the engine 10 is described with the low-speed compressor 44 being located in front of the high-speed compressor 34, in some embodiments, the compressors 34, 44 may be in an interleaved arrangement. Additionally or alternatively, although the engine 10 is described with the high-speed turbine 36 being located in front of the low-speed turbine 42, in some embodiments, the turbines 36, 42 may similarly be in an interleaved arrangement.

[0030] Still referring Figure 1 , the turbine 30 is generally enclosed in a shroud 48. Additionally, it can be understood that the shroud 48 at least partially defines an inlet 50 and an exhaust 52 and includes a turbomachinery flow passage 54 extending between the inlet 50 and the exhaust 52. The inlet 50 is a circular or axially symmetric 360-degree inlet 50 in the illustrated embodiment, which is located between the rotor blade assembly 12 and the stationary vane assembly 18 and provides a path for the incoming atmospheric air to enter the turbomachinery flow passage 54 (as well as the compressors 44, 34, the combustion section 40, and the turbines 36, 42) inside the guide vanes 28 along the radial direction R. Such a location may be advantageous for various reasons, including managing icing performance and protecting the inlet 50 from various objects and materials that may be encountered during operation.

[0031] However, in other embodiments, the inlet 50 may be located at any other suitable location, for example, behind the vane assembly 18, arranged in a non-axially symmetric manner, and so on.

[0032] As shown, the rotor assembly 12 is driven by the turbine 30, and more specifically, by the low-speed spool 55. More specifically, still Figure 1 the engine 10 in the illustrated embodiment includes a power gearbox 56, and the rotor assembly 12 is driven by the low-speed spool 55 of the turbine 30 passing through the power gearbox 56. In this way, the rotating rotor blades 16 of the rotor assembly 12 can rotate about the axis 14 and generate thrust to push the engine 10 and thus push the associated aircraft in the forward direction F. For example, in some embodiments, one or more engines constructed in a manner similar to Figure 1 the exemplary engine 10 described in Figure 1 , 4 and / or 5 can be incorporated into and utilized in an aircraft.

[0033] The power gearbox 56 may include a gear set for reducing the rotational speed of the low-speed spool 55 relative to the low-speed turbine 42, such that the rotor assembly 12 can rotate at a rotational speed slower than that of the low-speed spool 55.

[0034] As described above, the engine 10 includes a vane assembly 18. The vane assembly 18 extends from the shroud 48 and is located behind the rotor assembly 12. The vanes 20 of the vane assembly 18 may be mounted on a fixed frame or other mounting structure and do not rotate relative to the central axis 14. For reference purposes, Figure 1 the forward direction is also depicted by an arrow F, which in turn defines the front and rear of the system. As Figure 1 shown, the rotor assembly 12 is located in front of the turbine 30 in a "pulling" configuration, and the exhaust port 52 is located behind the guide vanes 28. As can be appreciated, the vanes 20 of the vane assembly 18 may be configured to straighten the airflow from the rotor assembly 12 (e.g., reduce vortices in the airflow) to improve the efficiency of the engine 10. For example, the size, shape, and configuration of the vanes 20 may be designed to create counteracting vortices in the airflow from the rotor blades 16 so that, in the downstream direction after the two rows of airfoils (e.g., blades 16, vanes 20), the degree of vorticity in the airflow is significantly reduced, which can translate into an increase in induced efficiency.

[0035] Still referring to Figure 1 , it may be desirable for the rotor blades 16, the vanes 20, or both to include pitch change mechanisms so that the airfoils (e.g., blades 16, vanes 20, etc.) can rotate independently or in combination relative to the pitch axis of rotation. Such pitch changes can be used to vary thrust and / or vortex effects under various operating conditions, including adjusting the magnitude or direction of the thrust generated on the rotor blades 16, or providing thrust reversal features, which may be useful under certain operating conditions, such as during aircraft landing, or desirably adjusting at least part of the acoustic noise generated by the rotor blades 16, the vanes 20, or the aerodynamic interaction of the rotor blades 16 relative to the vanes 20. More specifically, for Figure 1 the embodiment of

[0036] However, it will be understood that Figure 1 the exemplary single rotor ducted fan engine 10 depicted in

[0037] 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 turbofan engine, a turboprop engine, a turboprop engine, a turbojet engine, etc. Further, for example, although the engine is described as a single non-ducted rotor engine, in other embodiments, the engine may include a multi-stage open rotor configuration, and various aspects of the disclosure described below may be incorporated herein.

[0038] Further, still in other exemplary embodiments, the engine 10 may be configured as a turbofan engine. For example, briefly referring to Figure 2 , an engine 10 according to another exemplary embodiment of the present disclosure is described. Figure 2 The exemplary embodiment of Figure 1 is constructed in substantially the same manner as the exemplary engine 10 described above with respect to Figure 1 , and the same or similar reference numerals may refer to the same or similar parts. However, as will be understood, for the illustrated embodiment, the engine 10 further includes a nacelle 80 that at least partially circumferentially surrounds the rotor assembly 12 and the turbine 30, defining a bypass channel 82 therebetween.

[0039] Now returning to reference Figure 1 , it should be further understood that the engine is integrated with the power system 100. The power system 100 generally includes an electric machine 102 coupled to at least one of a high voltage system (or core 32) or a low voltage system, an energy storage unit 104, and an auxiliary power unit 106 for the illustrated embodiment. The auxiliary power unit 106 may include an internal combustion engine driving a generator and may be remote from the engine 10. For example, in at least some exemplary embodiments, the auxiliary power unit 106 may be located within the fuselage of an aircraft utilizing the engine, such as at the rear end of the aircraft (see, for example, Figure 1 ).

[0040] Further, for the illustrated embodiment, the power system 100 includes a power bus 108 that electrically connects the various components of the power system 100. The power bus 108 may be, for example, one or more electrical wires arranged in any suitable configuration.

[0041] Further, for the illustrated embodiment, the electric machine 102 of the power system 100 is a low-speed (“LS”) electric machine 102A that is coupled to a low-pressure system of the engine. More specifically, for the illustrated embodiment, the LS electric machine 102A is embedded within the engine 10, at a location within or behind the turbine section of the engine 10, and internally along a radial direction R through the core airflow path 54 of the engine 10. However, it is understood that in other exemplary embodiments, the LS electric machine 102A may alternatively or additionally be configured in other suitable manners. For example, in other embodiments, the LS electric machine 102A may be embedded within the compressor section of the engine 10, may be located radially outside the core airflow path 54 (and, for example, within the cowling 48), may be driven by a gearbox (such as an accessory gearbox), and so forth.

[0042] In addition, for the illustrated embodiment, the LS electric machine 102A is not the only electric machine 102 of the power system 100 that is integrated with the engine 10. More specifically, the power system 100 further includes a high-speed (“HS”) electric machine 102B that is coupled to the high-pressure system / core of the engine 10 and is in electrical communication with the power bus 108. In the illustrated embodiment, the HS electric machine 102B is also embedded within the engine 10 at a location inside the core airflow path 54. However, in the illustrated embodiment, the HS electric machine 102B is located within the compressor section of the engine 10. It is understood that in other embodiments, the HS electric machine 102B may alternatively be positioned radially outside the core airflow path 54 and driven, for example, by a gear connection. For example, in certain embodiments, the HS electric machine 102B may be coupled to an accessory gearbox (not shown) that is in turn coupled to the high-pressure system of the engine 10.

[0043] In at least some exemplary embodiments, the energy storage unit 104 may include one or more batteries. Additionally, or alternatively, the energy storage unit 104 may include one or more supercapacitor arrays, one or more ultracapacitor arrays, or both. In at least some embodiments, the energy storage unit 104 may be configured to provide at least 5 kilowatts (kW) of energy to the power system 100, such as at least 50 W, such as at least 50 W, such as at least 250 kW, such as at least 300 kW, such as at least 350 kW, such as at least 400 kW, such as at least 500 kW, such as up to 5 megawatts (MW), such as up to 10 megawatts (MW). In one or more of these configurations, the amount of power provided may refer to the peak power output at any instantaneous moment during the discharge period. Additionally, the energy storage unit 104 may be configured to provide such power for at least two minutes, such as at least three minutes, such as at least five minutes, such as up to one hour. For example, the energy storage unit 104 may be configured to store at least 10 kW-minutes of power, such as at least 50 kW-minutes of power, such as at least 250 kW-minutes of power, such as at least 500 kW-minutes of power, such as at least 1000 kW-minutes of power, such as at least 1500 kW-minutes of power, such as up to 250 MW-minutes of power. In one or more of these configurations, the amount of power provided above may be the peak power output at any instantaneous moment during the discharge period.

[0044] As Figure 1 shown in FIG. 1, the engine 10 includes one or more accessory systems. Specifically, for the illustrated embodiment, the engine 10 includes a blower 110 and a lubrication system 112. In the illustrated embodiment, the blower 110 is located within the under-cowl region (the region below the cowl 48 and outside the turbomachinery flow path 54) and is electrically connected to the energy storage unit. The blower 110 may be configured to provide an air flow to the compressor section, the combustion section, or both, from, for example, an ambient location (such as ambient air drawn in through a bypass duct), air drawn from the interior space of an aircraft wing, a location below the cowl 48, or a combination of these locations. The blower 110 may operate after the engine 10 is shut down to maintain an air flow through certain components of the engine 10, thereby preventing or minimizing the "back-diffusion" of heat within the components of the engine 10 into, for example, the fuel nozzles of the combustion section and heating such fuel nozzles to a temperature that would cause any remaining fuel to coke. For example, in some exemplary aspects, the blower may create an air flow through the turbomachinery flow path 54, the under-cowl region, or both. In particular, the blower 110 may provide an air flow from the under-cowl region through a compressor or combustion chamber port and through the compressor and combustion sections of the engine 10 into the turbomachinery flow path 54 (as Figure 1as shown by the dashed line). Optionally, the blower 110 may draw air flow from the turbomachinery flow path through the compressor and / or combustion chamber section of the engine 10 to provide desired ingestion protection.

[0045] In addition, as described above, for the illustrated embodiment, the engine 10 includes a lubrication system 112. The lubrication system 112 may be a lubrication system for a low-pressure system such that when driven by the power provided by the power system 100, the lubrication system 112 circulates lubricating fluid through various parts of the low-pressure system (e.g., bearings, reservoirs, heat exchangers, etc.). In this way, when, for example, the LS motor 102A rotates the low-pressure system when the engine is in an electric operation mode and the fuel flow to the combustion section has stopped, the lubrication system 112 can continue to support the lubrication function of the low-pressure system of the engine 10.

[0046] However, it should be understood that in another exemplary embodiment, the lubrication system 112 may additionally or alternatively include a lubrication system for a high-pressure system such that when driven by the power provided by the power system 100, the lubrication system 112 circulates lubricating fluid through various parts of the high-pressure system (e.g., bearings, reservoirs, heat exchangers, etc.).

[0047] Still referring to Figure 1 , the exemplary power system 100 is operably connected to a controller 116. The controller 116 may be an engine controller of the engine 10 (e.g., a full-authority digital engine control controller), may be an aircraft controller, may be a controller dedicated to the power system 100, and so on.

[0048] The controller 116 may be configured to receive data indicating various operating conditions and parameters of the engine 10 during engine operation. For example, as can be understood from Figure 1 the engine 10 includes one or more sensors 114 that are configured to sense data indicating various operating conditions and parameters of the engine, such as rotational speed, temperature, pressure, vibration, etc. For example, one or more sensors 114 may sense data indicating temperature parameters within the engine, such as exhaust gas temperature, combustion section temperature, compressor outlet temperature, etc. Additionally, or alternatively, one or more sensors 114 may sense data indicating the speed of the engine, such as the rotational speed of the low-pressure system, the rotational speed of the high-pressure system, the rotational speed of the rotor section 12, etc. It will be understood that although in Figure 1A single sensor 114 is depicted, but multiple sensors 114 providing the above functions may be located throughout the engine to sense relevant data. Additionally, as will be understood from the description herein, the controller 116 may also be configured to receive data from other sources, such as from an aircraft incorporating the engine, e.g., from one or more sensors of an aircraft incorporating the engine. In this manner, the controller 116 can receive data indicating the altitude of the aircraft, signals for engaging supplementary power, etc., from, for example, a pilot or other operator.

[0049] With particular reference to the operation of the controller 116, in at least some embodiments, the controller 116 may include one or more computing devices 118. The computing device 118 may include one or more processors 118A and one or more memory devices 118B. The one or more processors 118A may include any suitable processing device, such as a microprocessor, a microcontroller, an integrated circuit, a logic device, and / or other suitable processing devices. The one or more memory devices 118B may include one or more computer-readable media, including but not limited to non-transitory computer-readable media, RAM, ROM, hard disks, flash drives, and / or other memory devices.

[0050] The one or more memory devices 118B may store information accessible by the one or more processors 118A, including computer-readable instructions 118C executable by the one or more processors 118A. The instructions 118C may be any set of instructions that, when executed by the one or more processors 118A, cause the one or more processors 118A to perform operations. In some embodiments, the instructions 118C may be executed by the one or more processors 118A to cause the one or more processors 118A to perform operations, such as any operations and functions that the controller 116 and / or the computing device 118 is configured to perform, operations for operating the power system 100 (e.g., method 300) as described herein, and / or any other operations or functions of the one or more computing devices 118. The instructions 118C may be software written in any suitable programming language or may be implemented in hardware. Additionally, and / or alternatively, the commands 118C may be executed on the processor 118A in logically and / or virtually independent threads. The memory device 118B may further store data 118D accessible by the processor 118A. For example, the data 118D may include data indicating the power flow, data indicating the operating conditions of the engine 10 / aircraft, and / or any other data and / or information described herein.

[0051] The computing device 118 may also include a network interface 118E for communicating, for example, with other components of the engine 10, an aircraft incorporating the engine 10, the power system 100, and the like. For example, in the described embodiment, as described above, the engine 10 includes one or more sensors 114 for sensing data indicative of one or more parameters of the engine 10 and various accessory systems, and the power system 100 includes an energy storage unit 104, an LS motor 102A, an HS motor 102B, and an auxiliary power unit. The controller 116 is operably coupled to these components via, for example, the network interface 118E such that the controller 116 can receive data indicative of various operating parameters sensed by the one or more sensors 114 during operation, various operating conditions of the components, and the like, and further can provide commands to control the amount of electrical current in the power system 100 and other operating parameters of these systems, for example, in response to data sensed by the one or more sensors 114 and other conditions.

[0052] The network interface 118E may include any suitable components for communicating with one or more network interfaces, such as including a transmitter, a receiver, ports, a controller, an antenna, and / or other suitable components. For example, in the illustrated embodiment, the network interface 118E is configured to communicate wirelessly with these components via a wireless communication network (as shown by the dashed communication lines in Figure 1 ).

[0053] The techniques discussed herein refer to computer-based systems and actions taken by, and information sent to and from, computer-based systems. Those of ordinary skill in the art will recognize that the inherent flexibility of computer-based systems allows for a large number of possible configurations, combinations, and partitioning of tasks and functions among components. For example, the processing discussed herein can be implemented using a single computing device or a combination of multiple computing devices working together. Databases, memories, instructions, and applications can be implemented on a single system or distributed across multiple systems. The distributed components can operate sequentially or in parallel.

[0054] Now referring to Figure 3 , a flowchart of a method 300 for operating a hybrid propulsion system of an aircraft is provided. The method 300 can be used in conjunction with one or more of the exemplary embodiments described above with reference to Figure 1 and / or with one or more other exemplary embodiments described herein. For example, in at least some exemplary aspects, the hybrid propulsion system can include a gas turbine engine having a high-pressure system, a low-pressure system, an electric motor coupled to at least one of the high-pressure system or the low-pressure system, and an energy storage unit.

[0055] Method 300 includes, at (302), operating the electric machine as a generator to charge the energy storage unit during a flight operation of the aircraft. For example, operating the electric machine as a generator to charge the energy storage unit at (302) may include operating the electric machine as a generator to charge the energy storage unit during a cruise operation, a descent operation of the flight operation, a climb operation of the flight operation, etc. of the aircraft.

[0056] Method 300 further includes, at (304), switching the gas turbine engine to an electric operation mode during or after an aircraft landing operation. In some exemplary aspects, switching the gas turbine engine to the electric operation mode at (304) may include, at (306), cutting off the fuel flow to the gas turbine engine. Additionally, in some exemplary aspects, switching the gas turbine engine to the electric operation mode at (304) may include, at (308), switching the gas turbine engine to the electric operation mode after the aircraft landing operation. Once all the wheels of the aircraft are on the ground after the aircraft descends, the landing operation of the aircraft may be considered complete. Additionally, or alternatively, once the carrier ground speed drops below approximately 40 miles per hour, once in response to the position of the carrier (e.g., once the carrier leaves the runway), a command indicating completion of the landing operation is received from the pilot or operator, etc., then the landing operation of the aircraft may be considered complete.

[0057] Method 300 further includes using the electric power from the energy storage unit in the electric operation mode to drive a system of the gas turbine engine to provide or assist in providing ground operation of the aircraft. More specifically, method 300 further includes, at (310), utilizing the electric power from the energy storage unit in the electric operation mode to drive a low-pressure system, a high-pressure system, or both, to provide or assist in providing ground operation of the aircraft. The term "ground operation" generally refers to any operation from when the aircraft lands (i.e., the wheels touch the ground) to when all engine operations stop (including any accessory systems that operate after the engine stops rotating and receives fuel). As will be understood from the discussion herein, ground operations may include taxiing the aircraft, rotating one or more aspects of the engine to prevent or mitigate rotor bowing or engine soakback, driving the accessory systems of the engine, etc.

[0058] More specifically, for the depicted exemplary aspects, at (310), operating the low voltage system, the high voltage system, or both using power from an energy storage unit in an electrical operating mode to provide or assist in providing ground operation of the aircraft includes, at (311), using an electric machine to drive the low voltage system, the high voltage system, or both using power from the energy storage unit in an electrical operating mode to generate thrust for the aircraft using a gas turbine engine to taxi the aircraft. More specifically, for the depicted exemplary aspects, at (311), using an electric machine to drive the low voltage system, the high voltage system, or both using power from the energy storage unit to generate thrust for the aircraft to taxi the aircraft includes, at (312), using the electric machine to drive the low voltage system. It is noted that driving the low voltage system using the electric machine at (312) can correspondingly rotate the rotor assembly of the engine (e.g., fan assembly, propeller, etc., e.g., Figure 1 the rotor assembly 12 of the exemplary engine 10) to generate thrust for the aircraft to taxi the aircraft.

[0059] In accordance with the above, it should be understood that in at least some exemplary embodiments, to provide a desired amount of thrust for taxiing the aircraft, driving the low voltage system using the electric machine at (312) can include, at (314), using the electric machine to provide at least about 200 horsepower (hp) to the gas turbine engine to generate thrust for the aircraft. For example, in some exemplary aspects, driving the low voltage system using the electric machine at (312) can include using the electric machine to provide at least about 250 horsepower, such as at least about 300 horsepower, such as up to about 1000 horsepower to the gas turbine engine. Additionally, in at least some exemplary aspects, driving the low voltage system using the electric machine at (312) can include, at (316), using the electric machine to provide at least about 200 horsepower to the gas turbine engine to generate thrust for the aircraft for at least about three minutes, such as at least about five minutes, such as at least about eight minutes, such as at least about ten minutes.

[0060] Still referring to Figure 3Exemplary aspects of the method 200 depicted, it should be understood that the exemplary method 300 can further be used to reduce the negative impact of thermal soakback within a gas turbine engine after the gas turbine engine has been shut down. For example, the depicted exemplary method 300 further includes, at (318), parking the aircraft. Parking the aircraft at (318) generally refers to placing the aircraft in a situation where it does not move for an extended period of time. As will be recognized, once the aircraft is parked, the airflow through the gas turbine engine of the aircraft can be reduced, which may create an opportunity for the residual heat within the engine to transfer to certain components, potentially damaging those components. For example, the residual heat can transfer from various rotating and core components of the high-pressure system and the combustion section to the fuel nozzles in the combustion section, and to accessories such as control units and other electronic devices. Additionally or alternatively, the residual heat can transfer to the fuel at locations other than the fuel nozzles, potentially heating the fuel near or within the gas turbine engine and causing coke formation.

[0061] Accordingly, for Figure 3 exemplary aspects of the method 300 shown, at (310), driving a low-pressure system, a high-pressure system, or both using an electric machine that uses electrical power from an energy storage unit in an electric operating mode to provide or assist in providing ground operation of the aircraft further includes, at (322), using the electric machine to drive the low-pressure system of the aircraft to mitigate engine soakback. It is noted that in at least some exemplary aspects, using the electric machine at (322) to drive the low-pressure system of the aircraft to mitigate engine soakback can include, at (323), using the electric machine to drive the low-pressure system of the aircraft to mitigate engine soakback after parking the aircraft at (318).

[0062] In this manner, using the electric machine at (322) to drive the low-pressure system of the aircraft can include inducing an airflow through the core airflow path of the engine by providing an airflow through the core, to reduce engine soakback. Additionally, or alternatively, using the electric machine at (322) to drive the low-pressure system of the aircraft can include inducing an airflow through the core airflow path of the engine by providing an airflow that causes the core to rotate, to reduce engine soakback.

[0063] For example, using the electric machine at (322) to drive the low-pressure system of the aircraft to mitigate engine soakback further includes, at (324), using the electric machine to rotate the low-pressure system at a rotational speed less than 200 revolutions per minute and greater than 1 revolution per minute. For example, in some exemplary aspects, using the electric machine at (322) to drive the low-pressure system of the engine to mitigate engine soakback can include using the electric machine to rotate the low-pressure system of the engine at a rotational speed less than 100 revolutions per minute, such as at a rotational speed less than 50 revolutions per minute, such as at a rotational speed less than 25 revolutions per minute. Operating the engine in accordance with one or more of these exemplary aspects can provide a desired amount of soakback mitigation for the engine.

[0064] In addition, it will be understood that for the depicted embodiments, using the electric machine at (322) to drive a low-pressure system of the aircraft to mitigate engine soakback may include, at (325), using the electric machine to drive a low-pressure system of the aircraft to mitigate engine soakback in response to a sensed condition or in response to any other operating condition. For example, the sensed condition may be a temperature parameter of the engine (e.g., data indicative of exhaust gas temperature, compressor outlet temperature, etc.), a time parameter (e.g., time after engine shutdown or aircraft parking, etc.), or any other suitable sensed condition. The sensed condition may be a condition sensed by a controller from one or more engine sensors. The operating condition may include the sensed condition and, for example, a manual signal received from an operator.

[0065] In addition, it will be recognized that method 300 may further include using the electric machine to rotate a high-pressure system of the engine at a relatively low speed to reduce a bowed rotor condition. For example, method 300 may rotate the high-pressure system at a rotational speed of less than about 50 revolutions per minute, such as less than 25 revolutions per minute, such as less than 10 revolutions per minute, such as less than 1 revolution per minute. Such a processing step may allow thermal soakback to be evenly distributed circumferentially about one or more rotors of, for example, the high-pressure system, potentially reducing or eliminating a rotor bow condition.

[0066] However, it will be understood that depending on, for example, certain flight operating conditions and engine operating conditions during a prior flight, and / or certain environmental conditions, such efforts may not be sufficient to mitigate engine soakback to the desired extent. For the depicted exemplary aspects, driving the low-pressure system, the high-pressure system, or both at (310) further includes, at (326), receiving data indicating that an engine temperature parameter exceeds a predetermined threshold, while using the electric machine at (322) to drive a low-pressure system of the aircraft to mitigate engine soakback. Receiving the data at (326) may include receiving the data from one or more engine sensors using a controller.

[0067] In response, driving the low-pressure system, the high-pressure system, or both at (310) further includes, at (328), using a second electric machine to drive a high-pressure system of the gas turbine engine to increase cooling of the gas turbine engine (in response to receiving the data at (326)). In at least some exemplary aspects, using the second electric machine at (328) to drive a high-pressure system of the gas turbine engine may include using the second electric machine to rotate the high-pressure system of the gas turbine engine at a rotational speed of less than 200 revolutions per minute and greater than 1 revolution per minute.

[0068] However, it should be noted that in other exemplary aspects, instead of receiving the engine temperature parameter at (326), method 300 may receive data indicative of an operating condition while using an electric motor to drive a low-pressure system of the aircraft to mitigate engine soakback, and method 300 may further include using a second electric motor to drive a high-pressure system of the gas turbine engine to increase cooling of the gas turbine engine in response to receiving data indicative of an operating condition. The operating condition may be a time parameter, a temperature parameter, a manual signal, or a combination thereof.

[0069] In addition, it will be understood that for the depicted exemplary aspects, using an electric motor to drive a low-pressure system of the gas turbine engine at (322) to mitigate engine soakback may include driving the low-pressure system of the electric motor for a period of time between at least about two minutes and two hours after parking the aircraft at (318), such as equal to at least five minutes, such as equal to at least ten minutes, such as up to 90 minutes, such as up to 1 hour.

[0070] Now referring Figure 4 , other exemplary aspects of method 300 are described. Figure 4 The method 300 shown in Figure 3 may be constructed in substantially the same manner as the method 300 depicted in Figure 4 . However, for the exemplary aspects of , method 300 further includes, at (330), providing power from an energy storage unit to an accessory system of the gas turbine engine after switching the gas turbine engine to an electric operation mode. In at least some exemplary aspects, providing power from an energy storage unit to an accessory system of the gas turbine engine at (330) may include providing power to an accessory engine cooling system (such as a core soakback blower), providing power to a lubrication system (such as a lubrication system for the low-pressure system, a lubrication system for the high-pressure system, or both), and the like. In these examples, the accessory system may alternatively, selectively, or additionally be configured to operate during or after engine shutdown, for example, once the aircraft reaches the gate and / or is parked. For example, an electric blower located in the lower cowl area and operating during or after engine shutdown may be enabled and operated for a predetermined amount of time based on relevant operating or environmental parameters (such as taxi time, ambient air temperature) to limit the peak soakback temperature (such as 350 degrees Fahrenheit to avoid fuel coking, or 200 degrees Fahrenheit to avoid damaging the electric motor (if provided) near the low-pressure turbine), or the blower may operate for a period of time based on a sensed temperature located, for example, at or near the burner nozzle, the T3 location, or the rear end of the low-pressure turbine. The blower may be configured to generate a forced air flow through the core (such as through a bleed air port, a plenum chamber, etc. located at the downstream end of the compressor) by drawing air from a third flow of the engine, a supercharger bypass valve, or through the lower cowl area (drawing air through the rear end of the engine).

[0071] Additionally, although not shown, it should be understood that in still some exemplary aspects, method 300 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, and the like.

[0072] Furthermore, in one or more of these exemplary aspects, the energy storage unit may be insufficient to meet all of the power requirements of the gas turbine engine, the electric machine, and / or the aircraft. In such an exemplary aspect, method 300 further includes, at (332), determining that the power requirements of the electric machine, the gas turbine engine, the aircraft, or a combination thereof exceed the available power of the energy storage unit after switching the gas turbine engine to the electric operation mode, and at (334), operating an auxiliary power unit of the aircraft to generate additional electrical power for the electric machine, the gas turbine engine, the aircraft, or a combination thereof.

[0073] In at least some exemplary aspects, determining the power requirements of the electric machine, the gas turbine engine, the aircraft, or a combination thereof at (332) may include determining that the power requirements of the electric machine exceed the available power of the energy storage unit of the electric machine after switching the gas turbine engine to the electric operation mode. Additionally, determining the power requirements may include, for example, estimating the remaining time of the low-pressure system of the electric machine driving the gas turbine engine to, for example, mitigate engine soak-back, 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, or the high-pressure system of the gas turbine engine, or both, to mitigate engine soak-back; receiving temperature data of the electric machine that indicates the need to operate an additional aircraft cooling system (such as an engine soak-back blower, a second electric machine rotating the high-pressure system, etc.); and so on.

[0074] Furthermore, operating the auxiliary power unit of the aircraft at (334) to generate additional electrical power for the electric machine, the gas turbine engine, the aircraft, or a combination thereof may include providing power from the auxiliary power unit to the energy storage unit and / or the electric machine via a power bus of the power system.

[0075] Furthermore, it will still be recognized that, as described above, in the electric operation mode, the ground operation described herein is powered by an electric machine driving the low-pressure system, the high-pressure system, or both with power from the energy storage unit. In at least some exemplary aspects, most of all of the power (“total power”) used by the electric machine while providing or assisting in providing ground operation may come from the electrical energy storage unit. For example, at least 25% of the total power may come from the electrical energy storage unit, such as at least about 35%, such as at least about 50%, such as at least about 65%, such as at least about 75%, such as at least about 90% of the total power may come from the electrical energy storage unit.

[0076] However, in some exemplary aspects, not all of the total power may come from the electrical energy storage unit. For example, in some exemplary aspects, up to 85% of the total power may come from the electrical energy storage unit, such as up to 75% of the total power may come from the electrical energy storage unit. The remainder of the total power may be provided from the APU and / or, depending on ground operations, from a power source external to the aircraft (e.g., ground power).

[0077] For example, in some exemplary aspects, when ground operations include taxiing the aircraft, the energy storage unit may provide all the power for taxiing the aircraft and may then provide supplementary power to rotate the engine to relieve rotor bending / sag, or vice versa. Optionally, supplementary power may be provided for peak output power demands (e.g., during taxiing) but not for lower power output demands (e.g., rotating the engine to relieve rotor bending / sag).

[0078] This written description uses examples to disclose the invention, including the best mode, and also enables those skilled in the art to practice the invention, including making and using any device or system and performing any combined method. The patentable scope of the invention is defined by the claims and may include other examples that occur to those skilled in the art. If these other examples include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements that do not differ substantially from the literal language of the claims, then these other examples are intended to be within the scope of the claims.

[0079] Other aspects of the invention are provided by the subject matter of the following clauses:

[0080] A method for operating a hybrid propulsion system of an aircraft, the hybrid propulsion system including a gas turbine engine having a high-pressure system, a low-pressure system, an electric machine coupled to at least one of the high-pressure system or the low-pressure system, and an energy storage unit, the method including: during flight operations of the aircraft, operating the electric machine as a generator to charge the energy storage unit; during or after landing operations of the aircraft, switching the gas turbine engine to an electric operation mode; and in the electric operation mode, driving the systems of the gas turbine engine using power from the energy storage unit through the electric machine to provide or assist in providing ground operations of the aircraft.

[0081] A method according to one or more of these clauses, wherein the systems for driving the gas turbine engine include driving the low-pressure system, the high-pressure system, or both, and wherein the ground operations include taxiing the aircraft.

[0082] A method according to one or more of these clauses, wherein driving the low-voltage system, the high-voltage system, or both by the electric machine includes, in the electric operation mode, driving the low-voltage system, the high-voltage system, or both by the electric machine using electric power from the energy storage unit to generate thrust sufficient to taxi the aircraft.

[0083] A method according to one or more of these clauses, wherein driving the low-voltage system, the high-voltage system, or both by the electric machine includes driving the low-voltage system by the electric machine.

[0084] A method according to one or more of these clauses, wherein driving the low-voltage system by the electric machine includes supplying at least about 200 horsepower to the gas turbine engine by the electric machine to generate thrust for the aircraft.

[0085] A method according to one or more of these clauses, wherein driving the low-voltage system by the electric machine includes supplying at least about 250 horsepower to the gas turbine engine by the electric machine to generate thrust for the aircraft, up to about 1000 horsepower.

[0086] A method according to one or more of these clauses, wherein driving the low-voltage system by the electric machine includes supplying at least about 200 horsepower to the gas turbine engine by the electric machine to generate thrust for the aircraft for at least about 5 minutes.

[0087] A method according to one or more of these clauses, wherein switching the gas turbine engine to the electric operation mode during or after landing of the aircraft includes cutting off the fuel flow to the gas turbine engine.

[0088] A method according to one or more of these clauses, further comprising: after switching the gas turbine engine to the electric operation mode, determining that the power demand of the electric machine, the aircraft, or both exceeds the available power from the energy storage unit; and operating an auxiliary power unit of the aircraft to generate additional electric power for the electric machine, the aircraft, or both.

[0089] A method according to one or more of these clauses, wherein the system driving the gas turbine engine includes driving the low-voltage system, the high-voltage system, or both, wherein the ground operation includes driving the low-voltage system of the aircraft by the electric machine to mitigate engine soakback, and wherein the method further comprises: parking the aircraft, wherein driving the low-voltage system of the aircraft by the electric machine to mitigate engine soakback includes, after parking the aircraft, driving the low-voltage system of the aircraft by the electric machine to mitigate engine soakback.

[0090] A method according to one or more of these clauses further comprises: receiving data indicating that an engine temperature parameter exceeds a predetermined threshold, while driving the low-pressure system of the aircraft by the motor to mitigate engine re-ingestion; and driving the high-pressure system of the gas turbine engine by a second motor to increase cooling of the gas turbine engine.

[0091] A method according to one or more of these clauses, wherein driving the low-pressure system of the aircraft by the motor to mitigate engine re-ingestion comprises rotating the low-pressure system at a rotational speed lower than 200 revolutions per minute and higher than 1 revolution per minute.

[0092] A method according to one or more of these clauses, wherein driving the system of the gas turbine engine comprises driving the low-pressure system, the high-pressure system, or both by the motor using electrical power from the energy storage unit in the electric operation mode to generate thrust through the gas turbine engine.

[0093] A method according to one or more of these clauses further comprises providing electrical power for the hybrid propulsion system from an auxiliary power unit, from a power source external to the aircraft, or both.

[0094] A method according to one or more of these clauses, wherein the energy storage unit comprises one or more batteries, a supercapacitor array, a supercapacitor bank, or a combination thereof.

[0095] A method according to one or more of these clauses, wherein the system is an accessory system of the gas turbine engine, and wherein the accessory system comprises an accessory engine cooling system, a lubrication system of the low-pressure system, or a core re-ingestion blower.

[0096] A method according to one or more of these clauses, wherein the ground operation comprises driving the low-pressure system of the aircraft by the motor to mitigate engine re-ingestion, and wherein the method further comprises: receiving data indicating an operating condition while driving the low-pressure system of the aircraft by the motor to mitigate engine re-ingestion; and driving the high-pressure system of the gas turbine engine by a second motor to increase cooling of the gas turbine engine in response to receiving the data indicating the operating condition, wherein the operating condition is a time parameter, a temperature parameter, a manual signal, or a combination thereof.

[0097] A method according to one or more of these clauses, wherein driving the system of the gas turbine engine comprises operating a blower to generate an air flow through a turbomachinery flow path of the gas turbine engine, a lower cowl region of the gas turbine engine, or both.

[0098] A method according to one or more of these clauses, wherein the blower is positioned within the lower shroud area and is electrically connected to the energy storage unit.

[0099] A method for operating a hybrid propulsion system of an aircraft, the hybrid propulsion system including a gas turbine engine having a high-pressure system and a low-pressure system, an electric machine coupled to at least one of the high-pressure system or the low-pressure system, and an energy storage unit, the method including: during a flight operation of the aircraft, operating the electric machine as a generator to charge the energy storage unit; during or after a landing operation of the aircraft, switching the gas turbine engine to an electric operation mode; driving the low-pressure system, the high-pressure system, or both by the electric machine to generate thrust to assist the aircraft in taxiing.

[0100] A method according to one or more of these clauses, wherein driving the low-pressure system, the high-pressure system, or both by the electric machine to generate thrust to assist the aircraft in taxiing includes driving the low-pressure system by the electric machine to generate thrust to assist the aircraft in taxiing.

[0101] A method for operating a hybrid propulsion system of an aircraft, the hybrid propulsion system including a gas turbine engine having a high-pressure system and a low-pressure system, an electric machine coupled to at least one of the high-pressure system or the low-pressure system, and an energy storage unit, the method including: during a flight operation of the aircraft, operating the electric machine as a generator to charge the energy storage unit; during or after a landing operation of the aircraft, switching the gas turbine engine to an electric operation mode; in the electric operation mode, driving the low-pressure system, the high-pressure system, or both by the electric machine using electrical power from the energy storage unit; after switching the gas turbine engine to the electric operation mode, determining that the electrical power demand of the electric machine, the aircraft, or both exceeds the available electrical power from the energy storage unit; and operating an auxiliary power unit of the aircraft to generate additional electrical power for the electric machine, the aircraft, or both.

Claims

1. A hybrid propulsion system for an aircraft, characterized in that, The hybrid propulsion system includes: A gas turbine engine having a high-pressure system and a low-pressure system; An electric motor coupled to the high-pressure system or the low-pressure system; An energy storage unit configured to store electric power, the energy storage unit being electrically connected to the electric motor; and An auxiliary power unit electrically connected to the electric motor; Wherein the electric motor is configured to operate as a generator during the flight operation of the aircraft to charge the energy storage unit; Wherein the gas turbine engine is configured to switch to an electric operation mode during or after the landing operation of the aircraft; Wherein the low-pressure system, the high-pressure system, or both are configured to be driven by the electric power from the energy storage unit through the electric motor in the electric operation mode; and Wherein when the determined electric power demand of the electric motor, the aircraft, or both exceeds the available electric power from the energy storage unit, the auxiliary power unit is configured to generate additional electric power for the electric motor, the aircraft, or both.

2. The hybrid propulsion system according to claim 1, wherein Wherein the auxiliary power unit includes an internal combustion engine driving a generator.

3. The hybrid propulsion system according to claim 1, wherein Wherein the auxiliary power unit is located away from the gas turbine engine.

4. The hybrid propulsion system according to claim 1, characterized in that, Wherein the auxiliary power unit is located at the rear end of the aircraft.

5. The hybrid propulsion system according to claim 1, characterized in that, Wherein the electric motor includes a low-speed electric motor coupled to the low-pressure system of the gas turbine engine.

6. The hybrid propulsion system according to claim 5, wherein, Wherein the low-speed electric motor is located at the rear of the turbine section of the gas turbine engine.

7. The hybrid propulsion system according to claim 5, wherein Wherein the electric motor includes a high-speed electric motor coupled to the high-pressure system of the gas turbine engine.

8. The hybrid propulsion system according to claim 7, characterized in that, Wherein the high-speed electric motor is located within the compressor section of the gas turbine engine.

9. The hybrid propulsion system according to claim 1, characterized in that, Wherein the electric motor is configured to provide at least 200 horsepower to the low-pressure system of the gas turbine engine to generate thrust for the aircraft.

10. The hybrid propulsion system according to claim 1, characterized in that, Wherein the electric motor is configured to provide at least 200 horsepower to the low-pressure system of the gas turbine engine to generate thrust for the aircraft for at least five minutes.

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