Compressor stall mitigation

Abstract

A method for operating a hybrid-electric propulsion system for an aircraft, the hybrid-electric propulsion system including a gas turbine engine having a compressor and an electric motor coupled to the compressor, the method comprising sensing data indicative of pressure within the compressor of the gas turbine engine; determining, based at least in part on the sensed data indicative of the pressure within the compressor of the gas turbine engine, that a condition within the compressor is within a threshold of a stall limit for the compressor; and, in response to determining that the condition within the compressor is within the threshold of the stall limit for the compressor, using the electric motor to modify the torque of the compressor to reduce a risk of compressor stall.

Description

[0001] CROSS-REFERENCE TO RELATED APPLICATIONS

[0002] This application is a non-provisional application claiming the benefit of priority under 35 U.S.C. §119(e) to U.S. Provisional Application No. 63 / 072,573, filed on August 31, 2020, the entire contents of which are incorporated herein by reference. Technical Field

[0003] The present subject matter generally relates to hybrid-electric systems for aircraft and aircraft engines, and methods for operating such systems to mitigate compressor stall. Background Art

[0004] Conventional commercial aircraft typically include 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 on a respective one of the aircraft's wings, such as in an underwing mounting position, separate from the wings and fuselage.

[0005] Hybrid-electric propulsion systems are being developed to improve the efficiency of conventional commercial aircraft. Various hybrid-electric propulsion systems include an electric motor driven by one of the aircraft's engines. The inventors of the present disclosure have proposed various configurations and / or methods to improve currently known hybrid-electric propulsion systems, particularly systems and methods for mitigating compressor stall in hybrid-electric propulsion systems. Summary of the Invention

[0006] Aspects and advantages of the invention will be set forth in part in the following description, or may be obvious from the description, or may be learned through practice of the invention.

[0007] In one exemplary aspect of the present disclosure, a method for operating a hybrid-electric propulsion system for an aircraft is provided. The hybrid-electric propulsion system includes a gas turbine engine having a compressor and an electric motor coupled to the compressor. The method includes: sensing data indicating a pressure within the compressor of the gas turbine engine; determining, based at least in part on the sensed data indicating a pressure within the compressor of the gas turbine engine, that a condition within the compressor is within a threshold of a stall limit for the compressor; and in response to determining that the condition within the compressor is within the threshold of the stall limit for the compressor, using the electric motor to modify the torque of the compressor to reduce a risk of compressor stall.

[0008] 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 accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the present invention and, together with the description, serve to explain the principles of the present invention. BRIEF DESCRIPTION OF THE DRAWINGS

[0009] A full and enabling disclosure of the invention, including the best mode thereof, to one of ordinary skill in the art is set forth in the specification with reference to the accompanying drawings, in which:

[0010] Figure 1 is a schematic cross-sectional view of a gas turbine engine according to an exemplary embodiment of the present disclosure.

[0011] Figure 2 is a schematic cross-sectional view of a gas turbine engine according to another exemplary embodiment of the present disclosure.

[0012] Figure 3 is a schematic cross-sectional view of a compressor of a gas turbine engine according to an exemplary aspect of the present disclosure.

[0013] Figure 4 is a schematic diagram of a compressor stage of a gas turbine engine according to an exemplary aspect of the present disclosure.

[0014] Figure 5 is a schematic cross-sectional view of a compressor of a gas turbine engine according to another exemplary aspect of the present disclosure.

[0015] Figure 6 is a graph describing an operating line and a stall line of an engine according to an exemplary aspect of the present disclosure.

[0016] Figure 7 is a flow chart of a method for operating a hybrid-electric propulsion system for an aircraft. DETAILED DESCRIPTION

[0017] Reference will now be made in detail to present embodiments of the present 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 are used to refer to like or similar parts of the invention.

[0018] As used herein, "exemplary" means "serving as an example, instance, or illustration." Any embodiment described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other embodiments. Furthermore, all embodiments described herein are to be considered exemplary unless otherwise indicated.

[0019] As used herein, the terms “first,” “second,” and “third” may be used interchangeably to distinguish one component from another, and not for the purpose of indicating the position or importance of each component.

[0020] The terms "forward" and "rearward" refer to relative positions within a gas turbine engine or vehicle and to the normal operating attitude of the gas turbine engine or vehicle. For example, with respect to a gas turbine engine, forward refers to a position near the engine's air intake, and rear refers to a position near the engine's nozzle or exhaust.

[0021] The terms "upstream" and "downstream" refer to the direction relative to the flow of a path. For example, with respect to fluid flow, "upstream" refers to the direction from which the fluid is flowing, while "downstream" refers to the direction toward which the fluid is flowing. However, as used herein, the terms "upstream" and "downstream" may also refer to the flow of electricity.

[0022] The singular forms "a," "an," and "the" include plural references unless the context clearly dictates otherwise.

[0023] As used herein throughout the specification and claims, approximating language is applied to modify any quantitative representation that can be permissibly varied without resulting in a change in the basic function to which it is related. Thus, a value modified by one or more terms, such as "about," "approximately," and "substantially," is not limited to the precise value specified. In at least some cases, approximating language can correspond to the precision of an instrument for measuring the value, or the precision of a method or machine for constructing or manufacturing a component and / or system. In at least some cases, approximating language can correspond to the precision of an instrument for measuring the value, or the precision of a method or machine for constructing or manufacturing a component and / or system. For example, approximating language can mean within 1, 2, 4, 5, 10, 15, or 20% of a single value, a range of values, and / or an endpoint of a range defining a value.

[0024] Here and throughout the specification and claims, range limitations are combined and interchanged, such ranges are identified and include all sub-ranges contained therein unless context or language indicates otherwise. For example, all ranges disclosed herein are inclusive of the endpoints, and the endpoints are independently combinable with each other.

[0025] In aircraft gas turbine engines, compressors are typically designed to operate within conservative, predetermined thresholds of compressor stall margin to reduce the risk of compressor stall under even the most severe operating conditions, assuming the engine undergoes an expected amount of degradation (e.g., wear of seals, minor damage, etc.), which could result in the compressor, and the entire engine, not being able to operate at or near its full potential.

[0026] In a hybrid-electric aviation gas turbine engine, an electric motor may be coupled to a compressor, for example, coupled to a shaft rotatable with the compressor. The present disclosure utilizes an electric motor to facilitate operation of the compressor closer to a compressor stall margin. More specifically, the present disclosure utilizes one or more sensors within a compressor section to sense data indicative of a stall condition within the compressor, or within a threshold indicating a stall condition within the compressor. In response, the electric motor may be utilized to add power / torque to the compressor so that the compressor and engine continue to operate at a desired level or in a desired manner without stalling the compressor or approaching a stall condition within the compressor.

[0027] Under certain conditions, a high frequency sensor may be used to sense data indicative of a compressor rotating stall condition. The motor may provide power to counter the compressor rotating stall condition and allow the compressor to continue operating at a desired level without experiencing compressor rotating stall.

[0028] Now refer to Figure 1 , provides 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 The exemplary gas turbine engine of FIG. 1 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. Figure 1 As can be seen, the engine 10 takes the form of an open rotor propulsion system and has a rotor assembly 12 that includes an array of airfoils, and more particularly an array of rotor blades 16, arranged about a central longitudinal axis 14 of the engine 10.

[0029] Moreover, as will be explained in more detail below, the engine 10 additionally includes a non-rotating bucket assembly 18 positioned rearward of the rotor assembly 12 (i.e., non-rotating relative to the central axis 14) that includes an array of airfoils also arranged about the central axis 14, and more specifically includes an array of buckets 20 arranged about the central axis 14.

[0030] Rotor blades 16 are arranged in a typically equally spaced relationship about centerline 14, with each blade having a root 22 and a tip 24 with a span defined therebetween. Similarly, buckets 20 are also arranged in a typically equally spaced relationship about centerline 14, with each bucket having a root 26 and a tip 28 with a span defined therebetween. Rotor assembly 12 also includes a hub 44 located forward of the plurality of rotor blades 16.

[0031] Furthermore, 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 appreciated 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. Furthermore, it will be appreciated that the terms "high" and "low" are used in the same context to distinguish between the two systems and are not meant to imply any absolute speed and / or pressure values.

[0032] The core 32 generally includes a high-speed compressor 34, a high-speed turbine 36, and a high-speed shaft 38 extending between 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 can 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 can include one or more configurations for receiving a mixture of fuel and air and providing a flow of combustion gases through the high-speed turbine 36 to drive the high-speed spool.

[0033] The low-speed system also includes a low-speed turbine 42, a low-speed compressor or supercharger 44, and a low-speed shaft 46 extending between 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.

[0034] Although engine 10 is described as having low-speed compressor 44 positioned ahead of high-speed compressor 34, in certain embodiments, compressors 34, 44 may be positioned in a staggered arrangement. Additionally or alternatively, although engine 10 is described as having high-speed turbine 36 positioned ahead of low-speed turbine 42, in certain embodiments, turbines 36, 42 may similarly be positioned in a staggered arrangement.

[0035] Still refer to Figure 1 , the turbine 30 is typically enclosed in a cowling 48. Furthermore, it will be appreciated that the cowling 48 at least partially defines an air inlet 50 and an exhaust 52, and includes a turbomachinery flow passage 54 extending between the air inlet 50 and the exhaust 52. The air inlet 50, in the illustrated embodiment, is an annular or axisymmetric 360-degree air inlet 50 that is located between the rotor blade assembly 12 and the stationary or stationary vane assembly 18 and provides a path for incoming atmospheric air to enter the turbomachinery flow passage 54 (and the compressor 44, 34, the combustion section 40, and the turbine 36, 42) inboard of the guide vanes 28 in a radial direction R, a location that may be advantageous for various reasons, including managing icing performance and protecting the air inlet 50 from various objects and materials that may be encountered during operation.

[0036] However, in other embodiments, the air inlet 50 may be positioned at any other suitable location, such as, for example, behind the bucket assembly 18 , arranged in a non-axisymmetric manner, etc.

[0037] 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 through the power gearbox 56. In this manner, the rotating rotor blades 16 of the rotor assembly 12 can rotate about the axis 14 and generate thrust to propel the engine 10, and thus the aircraft associated therewith, in a forward direction F.

[0038] 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 so that the rotor assembly 12 may rotate at a slower rotational speed than the low-speed spool 55 .

[0039] As mentioned above, the engine 10 includes the bucket assembly 18. The bucket assembly 18 extends from the fairing 48 and is located behind the rotor assembly 12. The buckets 20 of the bucket assembly 18 can be mounted on a stationary 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 arrow F, which in turn defines the front and rear of the system. Figure 1 As shown, the rotor assembly 12 is positioned in a "pull" configuration ahead of the turbine 30, with the exhaust port 52 positioned behind the guide vanes 28. As can be appreciated, the vanes 20 of the vane assembly 18 can be configured to straighten the airflow from the rotor assembly 12 (e.g., reduce swirl in the airflow) to improve the efficiency of the engine 10. For example, the size, shape, and configuration of the vanes 20 can be designed to impart a reactionary swirl to the airflow from the rotor blades 16 such that, in the downstream direction behind the two rows of airfoils (e.g., blades 16, vanes 20), the airflow has a significantly reduced degree of swirl, which can translate into an improvement in induced efficiency.

[0040] Still refer to Figure 1It may be desirable for rotor blades 16, vanes 20, or both to include pitch variation mechanisms whereby the airfoils (e.g., blades 16, vanes 20, etc.) may be rotated independently or in conjunction with one another relative to a pitch axis of rotation. Such pitch variation may be used to vary thrust and / or swirl effects under various operating conditions, including adjusting the magnitude or direction of thrust generated at rotor blades 16, or providing a thrust reversal feature, which may be useful under certain operating conditions, such as during landing of an aircraft, or to desirably adjust acoustic noise generated at least in part by rotor blades 16, vanes 20, or from the aerodynamic interaction of rotor blades 16 relative to vanes 20. More specifically, for Figure 1 In the embodiment of the present invention, rotor assembly 12 is described as having a pitch variation mechanism 58 for rotating rotor blades 16 about their respective pitch axes 60 , and bucket assembly 18 is described as having a pitch variation mechanism 62 for rotating buckets 20 about their respective pitch axes 64 .

[0041] However, it is understandable that Figure 1 The exemplary single-spool, non-ducted engine 10 depicted in FIG is provided for example only. In other exemplary embodiments, the engine 10 may have any other suitable configuration, including, for example, any other suitable number of shafts or spools, turbines, compressors, etc.; fixed pitch blades 16, 20, or both; a direct drive configuration (i.e., gearbox 56 may not be included); and the like. For example, in other exemplary embodiments, the engine 10 may be a three-spool engine with a medium-speed compressor and / or turbine. In such a configuration, it will be understood that the terms "high" and "low" herein with respect to turbine, compressor, or spool speeds and / or pressures are used for convenience in distinguishing between components, do not require any specific relative speeds and / or pressures, and do not exclude additional compressors, turbines, and / or spools or shafts.

[0042] 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 turboshaft engine, a turboprop engine, a turbojet engine, etc. Furthermore, 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.

[0043] Further, in still other exemplary embodiments, the engine 10 may be configured as a ducted turbofan engine. For example, referring briefly to Figure 2 , an engine 10 according to another exemplary embodiment of the present disclosure is described. Figure 2 The configuration of the exemplary embodiment is similar to that described above with respect to Figure 1The exemplary engine 10 of FIG. 1 is substantially identical to the exemplary engine 10 of FIG. 1 , and like or similar reference numerals may refer to like or similar parts. However, as will be appreciated, 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 passage 82 therebetween.

[0044] Now back Figure 1 As will be appreciated, 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, and an energy storage unit 104.

[0045] Further, for the illustrated embodiment, the electric machine 102 of the power system 100 is a low-voltage electric machine 102A coupled to the low-voltage system of the engine. More specifically, for the illustrated embodiment, the low-voltage electric machine 102A is embedded within the engine 10, located within or behind the turbine section of the engine 10, and inboard of the core airflow path 54 through the engine 10 in the radial direction R. However, it will be appreciated that in other exemplary embodiments, the low-voltage electric machine 102A may additionally or alternatively be configured in other suitable manners. For example, in other embodiments, the low-voltage electric machine 102A may be embedded within the compressor section of the engine 10, may be located outboard of the core airflow path 54 in the radial direction R (and, for example, within the cowling 48), etc.

[0046] Furthermore, for the embodiment shown, the LP motor 102A is not the only motor 102 of the electrical system 100 that is integrated with the engine 10. More specifically, the electrical system 100 further includes an HP motor 102B that is coupled to the high voltage system / core of the engine 10 and in electrical communication with the electrical bus 108. In the embodiment shown, the HP motor 102B is also embedded within the engine 10 at a location inboard of the core airflow path 54. However, in the embodiment shown, the HP motor 102B is located within the compressor section of the engine 10. It will be appreciated that in other embodiments, the HP motor 102B may alternatively be positioned outboard of the core airflow path 54 along the radial direction R and driven by, for example, a gear connection. For example, in some embodiments, the HP motor 102B may be coupled to an accessory gearbox (not shown), which in turn is coupled to the high voltage system of the engine 10.

[0047] 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 kilowatts, such as at least 50 kilowatts, such as at least 250 kilowatts, such as at least 300 kilowatts, such as at least 350 kilowatts, such as at least 400 kilowatts, such as at least 500 kilowatts, such as up to 5 megawatts (MW), such as up to 10 megawatts (MW). In addition, the energy storage unit 104 may be configured to provide such electricity for at least two minutes, such as at least three minutes, such as at least five minutes, such as up to one hour. In addition, in other embodiments, the energy storage unit 104 may be configured to provide such electricity for any other suitable duration.

[0048] Additionally, for the illustrated embodiment, power system 100 includes a power bus 108 that electrically connects the various components of power system 100. Power bus 108 may be, for example, one or more electrical wires arranged in any suitable configuration.

[0049] Still refer to Figure 1 Although not depicted in the exemplary embodiment, it is understood that the exemplary power system may also include an auxiliary power unit. The auxiliary power unit, if included, may include an internal combustion engine driving a generator and may be located remotely from the engine 10. For example, in at least some exemplary embodiments, the auxiliary power unit (if provided) may be located within the fuselage of the aircraft using the engine 10, for example, at the rear end of the aircraft, and electrically coupled to the power bus 108.

[0050] Still refer to Figure 1 , the exemplary power system 100 is operatively connected to a controller 116. The controller 116 may be an engine controller for the engine 10 (e.g., a full authority digital engine control controller), an aircraft controller, a controller dedicated to the power system 100, or the like.

[0051] The controller 116 may be configured to receive data indicative of various operating conditions and parameters of the engine 10 during operation of the engine 10. For example, the engine 10 includes one or more sensors 114 configured to sense data indicative of various operating conditions and parameters of the engine 10, such as rotational speed, temperature, pressure, vibration, etc. However, more specifically, for Figure 1In the exemplary embodiment described in , the one or more sensors 114 include a first speed sensor 114A configured to sense data indicative of one or more parameters of the rotor assembly 12 (e.g., rotational speed, acceleration, torque on a rotor shaft driving the rotor assembly 12, etc.); a second sensor 114B configured to sense data indicative of the compressor (e.g., pressure within the high-pressure compressor 34, pressure within the low-pressure compressor 44, etc.); a third sensor 114C configured to sense data indicative of one or more combustion section parameters (e.g., temperature within the combustion section 40, fuel flow to the combustion section 40, one or more pressures within or around the combustion section 40, etc.), one or more high-pressure turbine parameters (e.g., turbine inlet temperature, rotational speed of the high-pressure turbine 36, etc.), or both; a fourth sensor 114D operable to sense data indicative of one or more parameters of the low-pressure system (e.g., rotational speed of the low-pressure spool 55); and a fifth sensor 114E configured to sense data indicative of one or more variable geometry components (e.g., position of one or more variable inlet guide vanes, outlet guide vanes, rotor blades 16, guide vanes 20, etc.).

[0052] 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 devices 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 device. 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, a hard drive, a flash drive, and / or other memory devices.

[0053] One or more memory devices 118B may store information accessible to one or more processors 118A, including computer-readable instructions 118C executable by one or more processors 118A. Instructions 118C may be any set of instructions that, when executed by one or more processors 118A, cause one or more processors 118A to perform operations. In some embodiments, instructions 118C may be executed by one or more processors 118A to cause one or more processors 118A to perform operations, such as any operations and functions for which controller 116 and / or computing device 118 are configured, operations for operating power system 100 (e.g., method 300), as described herein, and / or any other operations or functions of one or more computing devices 118. Instructions 118C may be software written in any suitable programming language or implemented in hardware. Additionally, and / or alternatively, instructions 118C may be executed on processor 118A in logically and / or virtually independent threads. Memory device 118B may further store data 118D accessible by processor 118A. For example, data 118D may include data indicative of power flow, data indicative of engine 10 / aircraft operating conditions, and / or any other data and / or information described herein.

[0054] The computing device 118 may also include a network interface 118E for communicating with, for example, other components of the engine 10, an aircraft incorporating the engine 10, the power system 100, and the like. For example, in the depicted 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, a low-power motor 102A, a high-power motor 102B, and an auxiliary power unit. The controller 116 is operatively 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 can further provide commands to control the current flow and other operating parameters of the power system 100, for example, in response to the data and other conditions sensed by the one or more sensors 114.

[0055] The network interface 118E may include any suitable components for interfacing with one or more networks, including, for example, transmitters, receivers, ports, controllers, antennas, 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 (e.g., a wireless communication network). Figure 1 (shown by the dashed communication line in the middle).

[0056] The technology discussed herein refers to computer-based systems and actions taken by, and information sent to, and from, computer-based systems. Those skilled in the art will recognize that the inherent flexibility of computer-based systems allows for a wide variety of possible configurations, combinations, and divisions of tasks and functions between components. For example, the processing discussed herein can be implemented using a single computing device or a combination of multiple computing devices. Databases, memories, instructions, and applications can be implemented on a single system or distributed across multiple systems. Distributed components can operate sequentially or in parallel.

[0057] Now refer to Figure 3 , depicts a close-up view of a compressor 202 of a gas turbine engine 200 including an electric machine 204. The compressor 202 may be a high-pressure compressor, such as Figure 1 and 2 The high pressure compressor 34, and in this manner, it can be understood that the gas turbine engine 200 can be operated in a manner similar to that described above with reference to, for example Figure 1 and 2 The exemplary gas turbine engine 10 is configured in the manner described.

[0058] However, it is worth noting that in other embodiments, the systems and methods of the present disclosure may be applied to low pressure compressors, such as those described above with reference to FIG. Figure 1 and 2 Additionally, or alternatively, in other exemplary embodiments, the systems and methods of the present disclosure may be applied to fans, such as ducted fans, such as Figure 2 In these alternative embodiments, the sensors discussed below may be arranged on the outside of the low pressure compressor rotor blades or fan blades, respectively.

[0059] Still refer to Figure 3 The gas turbine engine 200 further includes an electric motor 204 coupled to the compressor 202, and the electric motor 204 may be similar to Figure 1 and 2 1B is configured as an exemplary motor 102B. However, alternatively, it may be configured in any other suitable manner. For example, for the embodiment shown, the motor 204 is embedded within the engine 200 (e.g., positioned inside the core air flow path 206 in the radial direction R). However, in other embodiments, the motor 204 may be connected via an offset connection (e.g., via a gear train extending through the core air flow path 206; see Figure 1 ) is coupled to the compressor 202. The electric machine 204 may be configured to add power to the compressor 202 through an axial shaft rotatable with the compressor 202 and / or extract power from the compressor 202 through a shaft rotatable with the compressor 202.

[0060] For the illustrated embodiment, the compressor 202 further includes a plurality of sensors 208 configured to sense data at a high frequency that is indicative of pressure within the compressor 202 and more specifically at one or more locations within the compressor 202, such as pressure at one or more locations within the core air flow path 206 of the compressor 202 along an axial direction A of the engine 200.

[0061] Additionally, or alternatively, reference is now also made to Figure 4 , a schematic diagram of a stage of compressor rotor blades 214 of the compressor 202 along an axial direction A is provided. It will be understood that, in certain exemplary embodiments, a plurality of sensors 208 may be configured to sense data at a plurality of locations along a circumferential direction C of the engine 200 within the core air flow path 206 of the compressor 202. Furthermore, it will be understood that the sensors 208 may be positioned above the blades 214, such as above the tips of the blades, and aligned with the blades along the axial direction A (each within a reference plane defined by the axial direction A and the radial direction R). In either configuration, the sensors 208 may be operable to sense data indicative of pressure fluctuations at the compressor element caused by blade tip vortices 216. For example, as the rotor blades 216 rotate in a first circumferential direction C1, vortices 216 may form behind one or more of the rotor blades 216, causing pressure fluctuations.

[0062] The frequency at which the sensor reads an operating characteristic associated with the compressor element (e.g., pressure fluctuations at the compressor element) can be approximately two thousand (2,000) times per second, or approximately equal to the update rate of one or more computing devices 210 (described below). In this manner, it is understood that, for example, in certain exemplary aspects, the term "high frequency" refers to a frequency of at least 500 Hz, such as at least 1000 Hz, such as at least 1500 Hz, and up to 10,000 Hz, such that a sensing loop closure occurs at least once every millisecond.

[0063] In certain exemplary embodiments, sensor 208 may be part of a compressor active stability management (CASM) system designed to protect the engine from compressor stalls. For example, the CASM system may receive data from sensor 208 and interpret the data in a manner that indicates an impending stall or that a stall margin remains.

[0064] Return to the specific Figure 3 , the sensor 208 is further configured to transmit such information to an engine controller 210 (e.g., a FADEC). The engine controller 210 may be configured similarly to the above with respect to Figure 1In this manner, it will be appreciated that the controller 210 is in operative communication with the motor 204 such that the engine controller 210 can provide instructions configured to regulate the power extracted from the compressor 202 by the motor 204 and / or the power added to the compressor 202 by the motor 204.

[0065] Notably, for the illustrated embodiment, the controller 210 communicates with the electric machine 204 via the power converter 112 , which may include any suitable power electronics, converters, etc. The systems described herein may provide power to the electric machine via a connection to the engine controller 210 , or alternatively, the engine controller 210 may instead control the operation of the power converter 112 (which is in electrical communication with a separate power source, such as a separate electric machine, an energy storage unit, an auxiliary power unit, etc.), which in turn allows the engine controller 210 to control the operation of the electric machine 204 .

[0066] In this manner, the system can determine the risk of stall within compressor 202. More specifically, utilizing high-frequency sensory data indicating the pressure within compressor 202 at specific locations within compressor 202, the system can determine how close compressor 202 is to a stall threshold (e.g., a rotating stall threshold). It will be appreciated that the term "rotating stall" generally refers to a localized interruption of airflow within compressor 202, which continues to provide compressed air, but with reduced efficiency. Rotating stall may occur when a small number of airfoils experience airfoil stall, disrupting localized airflow without destabilizing compressor 202. Stalled airfoils may create relatively stagnant pockets of air that, rather than moving in the direction of flow, rotate in a circumferential direction C of compressor 202. In certain exemplary embodiments, there may be only one "stalled" airfoil, but the rotating stall may grow from there, propagating to multiple airfoils, resulting in a stalled airfoil surge and a more pronounced compressor stall.

[0067] To combat this rotating stall, the system can modify the torque on compressor 202 (e.g., the torque on a shaft or spool drivingly coupled to compressor 202) to avoid the stall condition, or to avoid or reduce the rotating stall condition. For example, in response to sensing data indicating a condition within a predetermined range of a compressor rotating stall threshold, the system can modify the torque on compressor 202 using motor 204. For example, in response to sensing data indicating a condition within a predetermined range of a compressor rotating stall threshold, the system can add torque to compressor 202 using motor 204.

[0068] The system can continue to sense data indicating the pressure within compressor 202 at a specific location within compressor 202 at a high frequency, while modifying the torque on compressor 202 and increasing or decreasing the amount of torque added to or extracted from compressor 202 in response to the sensed data. In this manner, the control of the system can be referred to as feedback control loop control. Furthermore, the amount of torque modification can be proportional to the value of the sensed data and how close the sensed data indicates compressor 202 is to a rotating stall condition.

[0069] For example, in certain exemplary aspects, the system may continue to high-frequency sense data indicative of pressure within compressor 202 at the same location (eg, at the same stage of compressor rotor blades 214 ) while modifying torque to ensure that a particular compressor rotating stall is mitigated.

[0070] Additionally, or alternatively, in other exemplary aspects, the system may continue to high-frequency sense data indicative of pressure within compressor 202 at other locations (e.g., at a stage upstream or downstream of compressor rotor blades 214) while modifying the torque to ensure that compressor rotating stall does not progress forward or backward. For example, in certain engine configurations, the location with the highest probability of experiencing compressor rotating stall moves forward or backward based on the overall pressure ratio (OPR) of the compressor. For example, as the OPR increases, the location with the highest probability of experiencing compressor rotating stall may move forward, while as the OPR decreases, the location with the highest probability of experiencing compressor rotating stall may move backward. Alternatively, the location with the highest probability of experiencing compressor rotating stall may move in opposite directions based on the OPR.

[0071] More specifically, it will be appreciated that during transient conditions (e.g., acceleration), the critical level of stall of compressor 202 may shift from the front end of compressor 202 to the rear end of compressor 202. Accordingly, the system may simultaneously monitor multiple sensors 208 (e.g., CASM system sensors) and control torque input and frequency to manage the limiting level of compressor 202 at any particular time. In this manner, it will be appreciated that the system may select a torque curve (amount, frequency, etc.) to match the position of the limiting level of compressor 202 to combat any rotating stall that may occur at that level.

[0072] It will also be understood that rotating stall can present as pressure fluctuations that define a sinusoidal pattern. In this manner, the system can sense data indicating pressure fluctuations (including the pattern of the pressure fluctuations) at a location within compressor 202 using sensor 208, and can apply torque using motor 204 at a frequency and amplitude configured to oppose the pressure pattern (e.g., at a sinusoidal pattern that is 180 degrees out of phase with the pattern of the pressure fluctuations, or more generally, varying the phase, frequency, and / or amplitude of the torque variations to reduce the amplitude of the pressure variations using a feedback loop). This pressure can be a pressure above a threshold indicating potential rotating stall.

[0073] As will be further appreciated from the foregoing, utilizing an electrical power source in this manner to modulate radial velocity according to a harmonic curve based on pressure variations sensed in sensor 208 helps to reduce boundary layer separation and non-separation conditions that produce sensed pressure variations, thereby restoring or increasing the margin between transient operating line 164 and stall line 254 (see FIG. Figure 6 ).

[0074] With such a system, the engine can be operated closer to the stall margin limit, particularly during transients, allowing for faster acceleration, smaller core engines, and overall higher performance, as the system described herein can prevent or minimize stall conditions by increasing power in response to data indicating that the engine (or more particularly the compressor 202) is approaching a stall condition. Furthermore, with such a system, the engine 200 can be designed to operate more optimally using different bleed valve controls (e.g., closing the bleed valve more), different variable inlet guide vane schedules, etc., and utilizing the present disclosure to mitigate any conditions approaching a compressor stall.

[0075] It is understandable that Figure 3 The system described in the embodiment is only by way of example, and in other embodiments, any other suitable configuration may be provided. Figure 5 As shown, the system may not include or utilize engine controller 210 , but instead may use higher frequency power electronics within power converter 112 , for example, to control torque modification on compressor 202 .

[0076] As mentioned above, the system can sense data indicating compressor stall at a high frequency. Figure 3 In the embodiment of FIG, the response time of the motor 204 is limited by the refresh rate of the computing device 210. Figure 5In some embodiments, the motor 204 can provide a faster response time for the system by bypassing the engine controller 210 and responding more quickly to high-frequency data sensed by the sensor 208. For example, in one or more of these embodiments, the system can utilize the motor 204 to modify the torque of the compressor 202 within one (1) second (e.g., within 0.5 seconds, such as within 0.1 seconds, such as within 0.01 seconds, such as within 0.001 seconds, such as within 0.0005 seconds) of sensing data indicating that the compressor is within a compressor stall threshold.

[0077] Furthermore, it will be appreciated that while in the illustrated embodiment, a plurality of high-frequency sensors are used to sense pressure data to determine a stall condition (or near-stall condition), in other exemplary aspects, the system may additionally or alternatively use other gas turbine engine parameters, such as data from other sensors indicating the rotational speed of one or more components of the engine (e.g., the compressor), temperatures within the engine (e.g., compressor outlet temperature), torque of one or more components of the engine (e.g., the compressor), etc. Furthermore, while for the exemplary embodiment described, the torsional response is provided by the electric machine 204, in other exemplary embodiments, the torsional response from the electric machine may be incorporated without action, such as modifying one or more variable geometry components (e.g., variable guide vanes), fuel flow, etc.

[0078] Now refer to Figure 6 , a graph describing the application of the above-described system is provided. The graph plots the variable Phi (indicative of the stall line of the engine 200) against the compressor rotational speed ("X-axis" 252) on a "Y-axis" 250. It will be appreciated that a new engine with minimal deterioration will be able to perform transportation operations, such as acceleration, better than an engine with more deterioration. As used herein, "deterioration" refers to, for example, actual engine deterioration (e.g., worn seals, scratches and minor deformations on airfoils, etc.), power extraction (e.g., using an electric motor to pull power from the engine shaft), adverse environmental conditions, etc.

[0079] It will be appreciated that, in at least some exemplary embodiments, the variable Phi may refer to a calculation of the fuel-air ratio (FAR) divided by the stoichiometric fuel-air ratio (FARst). Alternatively, Phi may be used to describe a parameter that varies with FAR / FARst. The modified Phi adjusts for the effects of temperature on compressor airflow.

[0080] Figure 6 Depicted in Figure 3The exemplary engine 200 has a stall line 254 and a stall buffer line 256 (or the nominal apex of the stall line). Generally speaking, the engine is designed to operate below the stall buffer line 256 in all operating modes and throughout the engine's life cycle. Such a configuration requires that the engine, which is originally capable of operating at higher performance (e.g., at a higher rotational speed, higher pressure, etc.), must be throttled to accommodate the relatively deteriorated expected extreme operating conditions.

[0081] However, the above reference Figures 3 to 5 The described system may allow newer engines to be set to operate closer to the stall buffer line 256. The benefits of this will be explained in more detail below.

[0082] Still refer to Figure 6 , Figure 6 Further depicted are the nominal engine operating lines at steady state (line 258) and transient state (line 260), which for the illustrated embodiment is the acceleration state of the engine 200. Figure 6 The operating lines of the older, more deteriorated engine are shown in a steady state (line 262) and in a transient state (line 264), which for the embodiment shown is also the acceleration state of the older, more deteriorated engine.

[0083] As shown, while the new engine is able to operate under transient conditions without exceeding the stall buffer line 256, under the same transient conditions, the older, more deteriorated engine exceeds the stall buffer line 256. However, the above reference to e.g. Figures 3 to 5 With the described system, older, more degraded engines can be operated in such transient conditions without significant stall risk due to the stall damping provided. For example, when the older, more degraded engine begins to operate above stall buffer line 256, high-frequency sensor 208 sensing data indicative of pressure within compressor 202 may indicate operation within a predetermined threshold for compressor stall, such as compressor rotating stall, which may be above stall buffer line 256. In response, the system may modify the torque on compressor 202 using electric motor 204, thereby enabling compressor 202 to operate at or below stall buffer line 256 while still providing a desired performance output.

[0084] In this manner, a gas turbine engine 200 including the system of the present disclosure can provide improved performance over the life of the engine. Additionally, or alternatively, such a configuration can allow for the utilization of a smaller core within the gas turbine engine 200, providing higher performance for a given size of the gas turbine engine 200, and / or maintaining a desired acceleration margin for the gas turbine engine 200 over the life of the gas turbine engine 200.

[0085] Now refer to Figure 7 , a flow chart of a method 300 for operating a hybrid electric propulsion system for an aircraft is provided. The hybrid electric propulsion system may be operated in a manner similar to that described above with reference to Figures 1 to 5 One or more exemplary hybrid-electric propulsion systems are described. For example, a hybrid-electric propulsion system may include a gas turbine engine having a compressor and an electric motor coupled to the compressor.

[0086] As depicted, method 300 generally includes sensing data indicative of pressure within a compressor of a gas turbine engine at (302). For the exemplary aspects described, sensing data indicative of pressure within the compressor of the gas turbine engine at (302) includes sensing data indicative of pressure within the compressor of the gas turbine engine at a high frequency at (304). Further for the exemplary aspects described, sensing data indicative of pressure within the compressor of the gas turbine engine at (302) includes sensing data indicative of pressure within the compressor of the gas turbine engine at (306) using a plurality of sensors. In certain exemplary aspects, the plurality of sensors may be arranged in an axial direction of the engine, the plurality of sensors may be arranged in a circumferential direction of the engine, or both.

[0087] Still refer to Figure 7 The method 300 further includes, at (308), determining, based at least in part on sensed data indicating pressure within a compressor of the gas turbine engine, that conditions within the compressor are within a threshold of a stall limit for the compressor. For example, in the described aspect, determining, at (308), that conditions within the compressor are within a threshold of a stall limit for the compressor includes, at (310), determining that conditions within the compressor are within a threshold of a rotating stall limit for the compressor.

[0088] The method 300 also includes, at (312), in response to determining that conditions within the compressor are within a threshold of a stall limit for the compressor, modifying the torque of the compressor using the electric motor to reduce the risk of the compressor stalling. In certain exemplary aspects, modifying the torque of the compressor using the electric motor at (312) includes, at (314), adding torque to the compressor using the electric motor. It will be appreciated that the electric motor can be used for purposes other than the present method 300. Thus, for example, when modifying the torque at (312), if the electric motor was already adding torque to the compressor when the determination was made at (308), the method 300 can increase the amount of torque added in response to the determination at (308), can reduce the amount of power extracted from components rotatable with the compressor in response to the determination at (308), and so on.

[0089] Further, method 300 can be configured to provide a response relatively quickly to the determination made at 308. For example, in certain exemplary aspects, modifying the torque of the compressor using the electric motor at (312) includes modifying the torque of the compressor using the electric motor at (316) within an amount of time less than 0.5 seconds (e.g., within an amount of time less than 0.05 seconds, such as within an amount of time less than 0.001 seconds, such as within an amount of time less than 0.0005 seconds) from determining that conditions within the compressor are within a threshold of a stall limit for the compressor.

[0090] It is worth noting that the method 300 can be operated as a closed-loop control method. Figure 7 , exemplary aspects of the described method 300 further include sensing additional data indicative of pressure within a compressor of the gas turbine engine at (318) while modifying the torque of the compressor using an electric motor; and increasing or decreasing the amount of torque added to the compressor in response to the additional sensed data at (320).

[0091] In certain exemplary aspects, sensing data indicative of pressure within a compressor of a gas turbine engine at (302) may include sensing data indicative of pressure within the compressor of the gas turbine engine at a first stage of the compressor at (322), and wherein sensing additional data indicative of pressure within the compressor of the gas turbine engine at (318) may include sensing additional data indicative of pressure within the compressor of the gas turbine engine at (324).

[0092] Additionally, or alternatively, in other exemplary aspects, sensing data indicative of pressure within the compressor of the gas turbine engine at (302) may include sensing data indicative of pressure within the compressor of the gas turbine engine at a first stage of the compressor at (326), and sensing additional data indicative of pressure within the compressor of the gas turbine engine at (218) may include sensing additional data indicative of pressure within the compressor of the gas turbine engine at a second stage of the compressor at (328).

[0093] In such exemplary aspects, it will be appreciated that sensing data indicative of pressure within the compressor of the gas turbine engine at (302) may include sensing data indicative of pressure within the compressor of the gas turbine engine at (330) while the compressor defines a first overall pressure ratio, and sensing additional data indicative of pressure within the compressor of the gas turbine engine at (318) may include sensing additional data indicative of pressure within the compressor of the gas turbine engine at (332) while the compressor defines a second overall pressure ratio.

[0094] This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal language of the claims.

[0095] Further aspects of the invention are provided by the subject matter of the following clauses.

[0096] A method of operating a hybrid-electric propulsion system for an aircraft, the hybrid-electric propulsion system including a gas turbine engine having a compressor and an electric motor coupled to the compressor, the method comprising: sensing data indicative of a pressure within the compressor of the gas turbine engine; determining, based at least in part on sensed data indicative of the pressure within the compressor of the gas turbine engine, that a condition within the compressor is within a threshold of a stall limit for the compressor; and, in response to determining that the condition within the compressor is within the threshold of the stall limit for the compressor, using the electric motor to modify the torque of the compressor to reduce a risk of compressor stall.

[0097] The method of one or more of these clauses, wherein determining that conditions within the compressor are within the threshold of the stall limit of the compressor comprises determining that conditions within the compressor are within the threshold of the rotating stall limit of the compressor.

[0098] The method of one or more of these clauses, wherein using the electric machine to modify the torque of the compressor comprises using the electric machine to add torque to the compressor.

[0099] The method according to one or more of these clauses, wherein sensing data indicative of the pressure within the compressor of the gas turbine engine comprises sensing data indicative of the pressure within the compressor of the gas turbine engine at a high frequency.

[0100] A method according to one or more of these clauses, wherein using the electric motor to modify the torque of the compressor includes using the electric motor to modify the torque of the compressor within an amount of time that is less than 0.5 seconds from determining that conditions within the compressor are within the threshold of the stall limit of the compressor.

[0101] The method according to one or more of these clauses further includes: sensing additional data indicative of pressure within the compressor of the gas turbine engine while modifying the torque of the compressor using the electric machine; and increasing or decreasing the amount of torque added to the compressor in response to the additional sensed data.

[0102] A method according to one or more of these clauses, wherein sensing data indicative of the pressure within the compressor of the gas turbine engine includes sensing data indicative of the pressure within the compressor of the gas turbine engine at a first stage of the compressor, and wherein sensing additional data indicative of the pressure within the compressor of the gas turbine engine includes sensing additional data indicative of the pressure within the compressor of the gas turbine engine at the first stage of the compressor.

[0103] A method according to one or more of these clauses, wherein sensing data indicative of the pressure within the compressor of the gas turbine engine includes sensing data indicative of the pressure within the compressor of the gas turbine engine at a first stage of the compressor, and wherein sensing additional data indicative of the pressure within the compressor of the gas turbine engine includes sensing additional data indicative of the pressure within the compressor of the gas turbine engine at a second stage of the compressor.

[0104] A method according to one or more of these clauses, wherein sensing data indicative of the pressure within the compressor of the gas turbine engine includes sensing data indicative of the pressure within the compressor of the gas turbine engine while the compressor defines a first overall pressure ratio, and wherein sensing additional data indicative of the pressure within the compressor of the gas turbine engine includes sensing additional data indicative of the pressure within the compressor of the gas turbine engine while the compressor defines a second overall pressure ratio.

[0105] The method according to one or more of these clauses, wherein the compressor is a high-pressure compressor, and wherein the electric motor is drivingly coupled to the high-pressure compressor.

[0106] The method of one or more of these clauses, wherein sensing data indicative of the pressure within the compressor of the gas turbine engine comprises sensing data indicative of the pressure within the compressor of the gas turbine engine with a plurality of sensors.

[0107] The method according to one or more of these clauses, wherein the plurality of sensors are arranged in an axial direction of the engine.

[0108] The method according to one or more of these clauses, wherein the plurality of sensors are arranged in a circumferential direction of the engine.

[0109] A hybrid electric propulsion system for an aircraft, the hybrid electric propulsion system comprising: a gas turbine engine having a compressor, an electric motor coupled to the compressor, and a controller, the controller comprising a memory and one or more processors, the memory storing instructions that, when executed by the one or more processors, cause the system to: sense data indicating a pressure within the compressor of the gas turbine engine; determine, based at least in part on the sensed data indicating a pressure within the compressor of the gas turbine engine, that a condition within the compressor is within a threshold of a stall limit for the compressor; and, in response to determining that the condition within the compressor is within the threshold of the stall limit for the compressor, use the electric motor to modify the torque of the compressor to reduce a risk of compressor stall.

[0110] The system of one or more of these clauses, wherein determining that the condition within the compressor is within the threshold of the stall limit of the compressor includes determining that the condition within the compressor is within the threshold of the rotating stall limit of the compressor.

[0111] The system of one or more of these clauses, wherein modifying the torque of the compressor using the electric machine comprises adding torque to the compressor using the electric machine.

[0112] The system of one or more of these clauses, wherein sensing data indicative of the pressure within the compressor of the gas turbine engine comprises sensing data indicative of the pressure within the compressor of the gas turbine engine at a high frequency.

[0113] A system according to one or more of these clauses, wherein using the electric motor to modify the torque of the compressor includes using the electric motor to modify the torque of the compressor within an amount of time that is less than 0.5 seconds from determining that conditions within the compressor are within the threshold of the stall limit of the compressor.

[0114] The system according to one or more of these clauses further includes: sensing additional data indicative of the pressure within the compressor of the gas turbine engine while modifying the torque of the compressor using the electric motor; and increasing or decreasing the amount of torque added to the compressor in response to the additional sensed data.

[0115] The system of one or more of these clauses, wherein the compressor is a high pressure compressor, and wherein the electric motor is drivingly coupled to the high pressure compressor.

Claims

1. A method for operating a hybrid electric propulsion system for an aircraft, characterized in that The hybrid electric propulsion system includes a gas turbine engine having a compressor and an electric machine coupled to the compressor, the method comprising: sensing data indicative of pressure within the compressor of the gas turbine engine; determining a location of a condition within the compressor of the gas turbine engine based at least in part on sensed data indicative of the pressure within the compressor, wherein the condition is within a threshold range of a rotating stall limit of the compressor; selecting a torque curve including torque magnitude and torque frequency to match the condition at the location within the compressor; and In response to determining the location of the condition within the compressor, the torque of the compressor is modified using the electric machine according to a selected torque curve to reduce a risk of compressor stall.

2. The method according to claim 1, characterized in that in, Using the electric machine to modify the torque of the compressor includes using the electric machine to add torque to the compressor.

3. The method according to claim 1, characterized in that in, Sensing data indicative of the pressure within the compressor of the gas turbine engine includes sensing data indicative of the pressure within the compressor of the gas turbine engine at a high frequency.

4. The method according to claim 1, wherein in, Using the electric motor to modify the torque of the compressor includes using the electric motor to modify the torque of the compressor in an amount of time less than 0.5 seconds from determining the condition.

5. The method according to claim 1, characterized in that Further include sensing additional data indicative of the pressure within the compressor of the gas turbine engine while modifying the torque of the compressor using the electric machine; and The amount of torque added to the compressor is increased or decreased in response to the additional sensed data.

6. The method according to claim 5, characterized in that in, Sensing data indicative of the pressure within the compressor of the gas turbine engine includes sensing data indicative of the pressure within the compressor of the gas turbine engine at a first stage of the compressor, and wherein sensing additional data indicative of the pressure within the compressor of the gas turbine engine includes sensing additional data indicative of the pressure within the compressor of the gas turbine engine at the first stage of the compressor.

7. The method according to claim 5, characterized in that in, Sensing data indicative of the pressure within the compressor of the gas turbine engine includes sensing data indicative of the pressure within the compressor of the gas turbine engine at a first stage of the compressor, and wherein sensing additional data indicative of the pressure within the compressor of the gas turbine engine includes sensing additional data indicative of the pressure within the compressor of the gas turbine engine at a second stage of the compressor.

8. The method according to claim 7, characterized in that in, Sensing data indicative of the pressure within the compressor of the gas turbine engine includes sensing data indicative of the pressure within the compressor of the gas turbine engine while the compressor defines a first overall pressure ratio, and wherein sensing additional data indicative of the pressure within the compressor of the gas turbine engine includes sensing additional data indicative of the pressure within the compressor of the gas turbine engine while the compressor defines a second overall pressure ratio.

9. The method according to claim 1, characterized in that in, The compressor is a high-pressure compressor, and wherein the electric motor is drivingly coupled to the high-pressure compressor.

10. The method according to claim 1, characterized in that in, Sensing data indicative of the pressure within the compressor of the gas turbine engine includes sensing data indicative of the pressure within the compressor of the gas turbine engine with a plurality of sensors.

11. The method according to claim 10, characterized in that in, The plurality of sensors are arranged along an axial direction of the engine.

12. The method according to claim 10, characterized in that in, The plurality of sensors are arranged along a circumferential direction of the engine.

13. A hybrid electric propulsion system for an aircraft, characterized in that: The hybrid electric propulsion system comprises: A gas turbine engine having a compressor, an electric motor coupled to the compressor, and a controller comprising a memory and one or more processors, the memory storing instructions that, when executed by the one or more processors, cause the system to: sensing data indicative of pressure within the compressor of the gas turbine engine; determining a location of a condition within the compressor of the gas turbine engine based at least in part on sensed data indicative of the pressure within the compressor, wherein the condition is within a threshold range of a rotating stall limit of the compressor; selecting a torque curve including torque magnitude and torque frequency to match the condition at the location within the compressor; and In response to determining the location of the condition within the compressor, the torque of the compressor is modified using the electric machine according to a selected torque curve to reduce a risk of compressor stall.

14. The system according to claim 13, wherein: in, Using the electric machine to modify the torque of the compressor includes using the electric machine to add torque to the compressor.

15. The system according to claim 13, wherein: in, Sensing data indicative of the pressure within the compressor of the gas turbine engine includes sensing data indicative of the pressure within the compressor of the gas turbine engine at a high frequency.

16. The system according to claim 13, wherein: in, Using the electric motor to modify the torque of the compressor includes using the electric motor to modify the torque of the compressor in an amount of time less than 0.5 seconds from determining the condition.

17. The system according to claim 13, wherein: Further including: sensing additional data indicative of the pressure within the compressor of the gas turbine engine while modifying the torque of the compressor using the electric machine; as well as The amount of torque added to the compressor is increased or decreased in response to the additional sensed data.

18. The system according to claim 13, wherein: in, The compressor is a high-pressure compressor, and wherein the electric motor is drivingly coupled to the high-pressure compressor.

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