阅读说明：本技术 用于飞行器的容错性混合电动推进系统 (Fault tolerant hybrid electric propulsion system for aircraft ) 是由 罗伯特·查尔斯·弘 保罗·罗伯特·杰明 于 2019-08-20 设计创作，主要内容包括：提供了一种混合电动推进系统和方法。更具体地,本公开涉及用于飞行器的混合电动推进系统的控制系统,其被构造为响应于扭矩源(例如发动机)上的快速电负载变化而快速且自动地采取动作。还提供了用于操作飞行器的混合电动推进系统的方法。(A hybrid electric propulsion system and method are provided. More particularly, the present disclosure relates to a control system for a hybrid electric propulsion system for an aircraft that is configured to take action quickly and automatically in response to rapid electrical load changes on a torque source (e.g., an engine). A method for operating a hybrid electric propulsion system for an aircraft is also provided.)

1. A hybrid electric propulsion system for an aircraft, comprising:

an engine;

a first electric machine mechanically coupled with the engine and configured to generate electrical power when driven by the engine;

one or more electrical power consuming loads electrically coupled with the first electrical machine and configured to receive electrical power from the first electrical machine, the one or more electrical power consuming loads including a second electrical machine;

a propulsion assembly mechanically coupled with the second motor and configured to generate thrust when driven by the second motor;

one or more sensing devices located on the aircraft for detecting one or more performance indicators indicative of electrical loads on the engines;

one or more controllers communicatively coupled with the one or more sensors, the one or more controllers configured to:

receiving the one or more performance indicators indicative of the electrical load on the engine from the one or more sensing devices;

determining whether there is a change in electrical load based at least in part on the one or more performance indicators; and is

Generating a control action in response to the electrical load change if there is the electrical load change.

2. The hybrid electric propulsion system of claim 1, further comprising:

a fuel control device configured to selectively control fuel flow to the engine, the fuel control device communicatively coupled with the one or more controllers, and wherein if the electrical load change is a decrease in load on the engine, the one or more controllers, in generating the control action, are configured to:

activating the fuel control device to reduce the fuel flow to the engine.

3. The hybrid electric propulsion system of claim 1, further comprising:

an electric brake system including a resistor bank and an electric switching arrangement for selectively electrically coupling the first electric machine and the resistor bank, and wherein if the electrical load change is a decrease in load on the engine, the one or more controllers, in generating the control action, are configured to:

activating the electrical switching apparatus to electrically couple the first electric machine and the resistor bank to direct electrical power from the first electric machine to the resistor bank.

4. The hybrid electric propulsion system of claim 1, further comprising:

a physical braking system mechanically coupled with the engine, and wherein if the electrical load change is a decrease in load on the engine, the one or more controllers, in generating the control action, are configured to:

activating the physical braking system to reduce the torque output of the engine.

5. The hybrid electric propulsion system of claim 1, further comprising:

an energy storage device electrically coupled with the first electric machine and electrically coupled with the second electric machine, and wherein the energy storage device is configured to receive an amount of excess power from the first electric machine if the electrical load change is a decrease in load on the engine.

6. The hybrid electric propulsion system of claim 1, wherein if the electrical load change is a decrease in load on the engine, the one or more controllers are further configured to:

activating one or more variable geometry components of the engine to operate the engine in a less efficient manner.

7. The hybrid electric propulsion system of claim 1, wherein the first electric machine is a generator and the second electric machine is an electric motor electrically coupled with the generator, and wherein, if the change in electrical load is a decrease in load on the engine, the one or more controllers are further configured to:

controlling at least one of the electric motor and the one or more electrical consuming loads to operate at a less efficient or high power.

8. The hybrid electric propulsion system of claim 1, wherein the aircraft includes one or more accessory loads, and wherein the hybrid electric propulsion system further comprises:

an electrical switching arrangement for selectively electrically coupling the first electric machine and the one or more accessory loads, and wherein, if the electrical load change is a reduction in load on the engine, the one or more controllers are further configured to:

activating the electrical switching apparatus to electrically couple the first electric machine and the one or more accessory loads to direct power from the first electric machine to the one or more accessory loads.

9. The hybrid electric propulsion system of claim 1, wherein if the electrical load change is an increase in load on the engine, the one or more controllers are further configured to:

controlling an energy storage device to deliver power to the second electric machine during transient periods.

10. A method for operating a hybrid electric propulsion system for an aircraft, the method comprising:

receiving one or more performance indicators indicative of an electrical load on the engine from one or more sensing devices located on the aircraft, the engine mechanically coupled with an electric machine configured to generate electrical power when driven by the engine, the electric machine electrically coupled with one or more electrical power consuming devices, wherein at least one of the one or more electrical power consuming devices is mechanically coupled with a propulsion assembly;

determining whether there is a reduction in electrical load on the engine based at least in part on the one or more performance indicators; and

generating a control action in response to the electrical load reduction if there is the electrical load reduction on the engine, wherein generating the control action comprises reducing fuel flow to the engine.

Technical Field

The present subject matter relates generally to hybrid electric propulsion systems for aircraft and methods of operating the same. More particularly, the present subject matter relates to a control system for a hybrid electric propulsion system for an aircraft that is configured to detect and take action in response to rapid electrical load changes.

Background

Parallel hybrid electric propulsion systems for aircraft typically include an internal combustion engine or other mechanically driven power plant that drives an electric generator to produce electric power. The internal combustion engine may also drive a thrust source of the aircraft, such as a propeller. The electrical power generated by the generator is used to drive an additional source of thrust. For example, electrical power may be provided to an electric motor that uses the electrical power to drive an additional thrust source, such as a propeller on the other side of the aircraft.

Such systems inherently contain single failures that can produce more severe variations in aircraft thrust and thrust asymmetry than non-hybrid systems due to the electrical coupling. Furthermore, rapid electrical load changes may lead to significant and unsafe aircraft handling problems. In addition, rapid changes in engine torque may cause overspeed problems and the like.

Accordingly, a control system and method for a hybrid electric propulsion system for an aircraft that addresses one or more of the challenges described above would be useful.

Disclosure of Invention

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.

In one aspect, the present disclosure is directed to a hybrid electric propulsion system for an aircraft. The hybrid electric propulsion system includes an engine and a first electric machine mechanically coupled with the engine and configured to generate electric power when driven by the engine. The hybrid electric propulsion system also includes one or more electrical power consuming loads electrically coupled with the first electric machine and configured to receive electrical power from the first electric machine. The one or more power consuming loads include the second electrical machine, and possibly other power consuming loads. Further, the hybrid electric propulsion system includes a propulsion assembly mechanically coupled with the second motor and configured to generate thrust when driven by the second motor. Further, the hybrid electric propulsion system includes one or more sensing devices located on the aircraft for detecting one or more performance indicators indicative of the electrical load on the engine. Additionally, the hybrid electric propulsion system includes one or more controllers communicatively coupled with the one or more sensors. The one or more controllers are configured to receive one or more performance indicators indicative of an electrical load on the engine from the one or more sensing devices; determining whether there is a change in electrical load based at least in part on the one or more performance indicators; and if there is a rapid electrical load change, generating a control action in response to the electrical load change.

In another aspect, the present disclosure is directed to a method for operating a hybrid electric propulsion system for an aircraft. The method includes receiving, from one or more sensing devices located on the aircraft, one or more performance indicators indicative of an electrical load on an engine, the engine mechanically coupled with an electric machine configured to generate electrical power when driven by the engine, the electric machine electrically coupled with one or more electrical power consuming devices, wherein at least one of the one or more electrical power consuming devices is mechanically coupled with the propulsion assembly. The method also includes determining whether there is a reduction in electrical load on the engine based at least in part on the one or more performance indicators. The method further comprises the following steps: generating a control action in response to the electrical load reduction if there is an electrical load reduction on the engine, wherein generating the control action comprises reducing fuel flow to the engine.

In another aspect, the present disclosure is directed to a hybrid electric propulsion system for an aircraft. The hybrid electric propulsion system includes an engine. The hybrid electric propulsion system also includes a first propulsion assembly mechanically coupled with the engine and configured to generate thrust when driven by the engine. Further, the hybrid electric propulsion system includes a first electric machine mechanically coupled with the engine and configured to generate electric power when driven by the engine. Further, the hybrid electric propulsion system includes a second electric machine electrically coupled with the first electric machine and configured to receive electric power from the first electric machine. Further, the hybrid electric propulsion system includes a second propulsion assembly mechanically coupled with the second motor and configured to generate thrust when driven by the second motor. The hybrid electric propulsion system also includes one or more sensing devices located on the aircraft for detecting one or more performance indicators indicative of the electrical load on the engine. Further, the hybrid electric propulsion system includes one or more controllers communicatively coupled with the one or more sensors, the one or more controllers configured to: receiving one or more performance indicators indicative of an electrical load on the engine from one or more sensing devices; determining whether there is a change in electrical load based at least in part on the one or more performance indicators; and if there is an electrical load change, generating a control action in response to the electrical load change.

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 invention and together with the description, serve to explain the principles of the invention.

Drawings

A full and enabling disclosure of the present invention, including the best mode thereof, directed to one of ordinary skill in the art, is set forth in the specification, which makes reference to the appended figures, in which:

FIG. 1 provides a perspective schematic view of an exemplary aircraft having a parallel hybrid electric propulsion system in accordance with an exemplary embodiment of the present disclosure;

FIG. 2 provides a schematic illustration of an exemplary control system for a parallel hybrid electric propulsion system for the aircraft of FIG. 1;

FIG. 3 provides a graph depicting torque load on an engine as a function of time, according to an exemplary embodiment of the present disclosure;

FIG. 4 provides a schematic illustration of another exemplary hybrid electric propulsion system, according to an exemplary embodiment of the present disclosure; and

fig. 5 provides a flow chart of an exemplary method according to an exemplary aspect of the present disclosure.

Detailed Description

Reference now will be made in detail to embodiments of the invention, one or more examples of which are illustrated in the drawings. Each example is provided by way of explanation of the invention, not limitation of the invention. In fact, it will be apparent to those skilled in the art that various modifications and variations can be made in the present invention without departing from the scope or spirit of the invention. For instance, features illustrated or described as part of one embodiment, can be used with another embodiment to yield a still further embodiment. Thus, it is intended that the present invention cover the modifications and variations of this invention provided they come within the scope of the appended claims and their equivalents. Further, as used herein, an approximation term, such as "approximately", "substantially" or "approximately", is meant to be within ten percent (10%) of the error. Further, as used herein, the terms "first," "second," and "third" may be used interchangeably to distinguish one element from another, and are not intended to denote the position or importance of the various elements. The terms "upstream" and "downstream" refer to relative directions with respect to fluid flow in a fluid path. For example, "upstream" refers to the direction from which the fluid flows, and "downstream" refers to the direction to which the fluid flows.

In general, the present disclosure relates to a hybrid electric propulsion system and method thereof. More particularly, the present disclosure relates to a control system for a hybrid electric propulsion system for an aircraft that is configured to take action quickly and automatically in response to rapid electrical load changes on a torque source (e.g., an engine).

Fig. 1 provides a perspective schematic view of an exemplary aircraft 100 having a hybrid electric propulsion system 110 in accordance with an exemplary embodiment of the present disclosure. As shown, for this embodiment, the aircraft 100 is a fixed-wing aircraft. In other embodiments, aircraft 100 may be other suitable types of aircraft, such as a rotary aircraft, a vertical takeoff and landing aircraft, a tiltrotor aircraft, an airship, an unmanned aircraft, and the like. Aircraft 100 extends between first end 102 and second end 104, e.g., along longitudinal axis L. In the embodiment shown in FIG. 1, first end 102 is a forward end of aircraft 100 and second end 104 is an aft or rear end of aircraft 100. Aircraft 100 includes a fuselage 106 and a pair of wings 108, each wing 108 extending laterally outward from fuselage 106. The aircraft may include various control surfaces for controlling the propulsion and movement of the aircraft 100. Example control surfaces include elevators, rudders, ailerons, spoilers, flaps, slats, air brakes or decorative devices, and the like. Various actuators, servomotors, and other devices may be used to manipulate various control surfaces and variable geometry components of aircraft 100. Further, as described above, the aircraft 100 includes a hybrid electric propulsion system 110 for generating thrust. More specifically, for this embodiment, the hybrid electric propulsion system 110 is a parallel hybrid electric propulsion system.

As shown in fig. 1, hybrid electric propulsion system 110 includes an engine 112 mounted to one of wings 108 of aircraft 100. Engine 112 may be any suitable aero-mechanical torque source. For example, the engine 112 is a gas turbine engine in the illustrated embodiment. The gas turbine engine may be configured as a turboprop (as shown in FIG. 1), or other suitable type of gas turbine engine, such as a turbofan engine, a turbojet engine, a turboshaft engine, or the like. In an alternative embodiment, the engine 112 may be a piston-driven engine or some other type of internal combustion engine, such as a rocket engine.

A first electric machine 114 is mounted to one of the wings 108 and is mechanically coupled with the engine 112. The first electric machine 114 is configured to generate electric power when driven by the engine 112. To generate electrical power, as will be appreciated, the first electric machine converts rotational energy received from the output shaft of the engine 112 into electrical energy that may be delivered to various components of the hybrid electric propulsion system 110, as described more fully below. Thus, the first electric machine 114 may function as a generator. In some preferred embodiments, the first electric machine 114 may function as a generator or a motor, as the case may be. For example, in the event of a complete failure of the first electric machine 114, a clutch 115 (fig. 2) or similar feature may be provided to disengage the first electric machine 114 from the engine 112.

The hybrid electric propulsion system 110 also includes a first propulsion assembly 116, the first propulsion assembly 116 being mechanically coupled with the engine 112 and configured to generate thrust when driven by the engine 112. For this embodiment, the first propulsion assembly 116 is a propeller or a fan. The blades of the propeller may be adjusted in unison by a plurality of pitch angles, for example, by activation of an actuation mechanism. Pitching of the blades may cause the propeller assembly to produce more or less thrust. In some embodiments, the first propulsion assembly 116 is mechanically coupled to the engine 112 in parallel with the first electric machine 114, for example, to avoid a single fault failure of the system. In an alternative embodiment, the first propulsion assembly 116 is mechanically coupled to the engine 112 in series with the first electric machine 114. Some of the torque output from the engine 112 is directed to a first electric machine 114, e.g., for generating electricity, and some of the torque output is supplied to a first propulsion assembly 116, e.g., for propelling the aircraft 100.

The first power converter 118 is electrically coupled with the first electric machine 114. The first power converter 118 provides an electronic interface between the first motor 114 and an electrical bus 120 of the propulsion system 110. As one example, the first power converter 118 may be a rectifier configured to convert Alternating Current (AC) generated by the first electric machine 114 to Direct Current (DC). The first power converter 118 may be a passive system including a plurality of diodes or an active system including various processing devices, semiconductor switches, and other electronic components.

The energy storage device 122 is electrically coupled with the first power converter 118, and thus the first electric machine 114. Energy storage device 122 may be embodied as, for example, one or more superconducting energy storage devices, batteries, or battery packs. Energy storage device 122 may be mounted within fuselage 106 or at another suitable location. As will be explained in greater detail herein, the energy storage device 122 may receive electrical power from the first electrical machine 114, and in some cases, may supply electrical power stored therein to various components of the propulsion system 110, such as the first electrical machine 114, the second electrical machine 126, and other loads.

Second power converter 124 is electrically coupled with first power converter 118 and energy storage device 122. Thus, the second power converter 124 is also in electrical communication with the first electric machine 114. The second power converter 124 provides electronics to interface the second motor 126 with the power bus 120 of the propulsion system 110. As one example, the second power converter 124 may be an inverter configured to convert DC current flowing through the power bus 120 into AC current, for example, to control the speed or torque of the second electric machine 126.

The second electric machine 126 is electrically coupled with the first electric machine 114 and is configured to receive electrical power from the first electric machine 114 (e.g., directly, indirectly via the energy storage device 122, or both). Further, the second electric machine 126 is configured to convert the electrical power received from the first electric machine 114 into rotational energy, e.g., to rotate an output shaft of the second electric machine 126. Thus, the second motor 126 may function as a motor. In some embodiments, the second electric machine 126 may function as a motor or a generator, as the case may be. The second power converter 124 may control the amount of power delivered to the second motor 126, for example, to control the speed or torque output of the output shaft of the second motor 126.

The second propulsion assembly 128 is mechanically coupled with the second motor 126, for example, via a coupling of an output shaft of the second propulsion assembly 128 and a power gearbox. The second propulsion assembly 128 is configured to generate thrust when driven by the second motor 126. For this embodiment, the second propulsion assembly 128 is a propeller or a fan, similar to the first propulsion assembly 116. The blades of the propeller may be adjusted in unison by a plurality of pitch angles, for example, by activation of an actuation mechanism. Pitching of the blades may cause the propeller assembly to produce more or less thrust. In some embodiments, the hybrid electric propulsion system 110 may include a plurality of electric machines, each coupled with one or more propulsion assemblies powered by the first electric machine 114. For example, the hybrid electric propulsion system 110 may include a third electric machine electrically coupled with the first electric machine 114 and configured to receive electric power therefrom. The third motor may be mechanically coupled to the third propulsion assembly to generate thrust. Further, the hybrid electric propulsion system 110 may include more motors and propulsion assemblies, which are similarly configured to the third motor and the third propulsion assembly.

In some cases, the hybrid electric propulsion system 110 may experience thrust asymmetry or significant handling issues due to rapid electrical load changes, particularly during rapid electrical load drops or motor torque losses. For example, if the first electric machine 114 fails or goes offline due to a detected error or other fault, the second electric machine 126, which is drawing power from the first electric machine 114, will cease producing output torque to drive the second propulsion assembly 128, and the electrical load on the engine 112 will rapidly decrease or drop. Similarly, a sudden loss or failure of the second electric machine 126 will cause the load or reactive torque on the first electric machine 114 to rapidly drop to zero (0) and the electrical load on the engine 112 will rapidly decrease or drop. This may cause the first electric machine 114 and/or the engine 112 to overspeed. Further, in this case, the thrust generated by the second propulsion assembly 128 will drop rapidly and, due to the removed electrical load on the first electric machine 114 and ultimately the engine 112, the thrust generated by the first propulsion assembly 116 will increase rapidly due to the reaction torque on the first electric machine 114 dropping to zero (0) with the torque output of the engine 112 remaining unchanged. The result is therefore an asymmetric thrust. That is, the thrust increases rapidly on one side of the aircraft, while the thrust decreases rapidly on the other side. According to an exemplary aspect of the present disclosure, the hybrid electric propulsion system 110 includes a control system 130 (fig. 2), the control system 130 including features to quickly and automatically take action in the event of such rapid electrical load changes. In particular, the control system 130 of the hybrid electric propulsion system 110 is configured to quickly and automatically account for electrical load variations on the order of microseconds as the engine 112 accelerates or decelerates to match the torque load of the electrical system on the engine 112. The engine 112 may take several seconds to accelerate or decelerate. Accordingly, the control system 130 is configured to take rapid and automatic action during this transition period. An exemplary control system is provided below.

FIG. 2 provides a schematic illustration of an exemplary control system 130 for the hybrid electric propulsion system 110 of the aircraft 100 of FIG. 1. As shown, the control system 130 includes one or more sensing devices 132 located on the aircraft 100 for detecting one or more performance indicators indicative of the electrical load on the engines 112. For this embodiment, at least one sensing device 132 is positioned adjacent to the shaft 138, for example, to sense or measure a performance indicator between the engine 112 and the first electric machine 114, the shaft 138 mechanically coupling the engine 112 with the first electric machine 114. The performance indicator may be the output torque of the engine 112 (or the input torque to the first electric machine 114), the rotational speed of the shaft 138, or other suitable indicator. The performance indicators may be used to calculate, predict, or estimate the electrical load on the engine 112. Further, at least one sensing device 132 is positioned adjacent the shaft 140, for example, to sense or measure a performance indicator between the second motor 126 and the second propulsion assembly 128, the shaft 140 mechanically coupling the second motor 126 with the second propulsion assembly 128. The performance indicator may be the output torque of the second motor 126 (or input torque to the second propulsion assembly 128), the rotational speed of the shaft 140, or other suitable parameter.

Other sensing devices 132 may include current, voltage, or speed sensors configured to measure various parameters of the first motor 114 and/or the second motor 126. From these performance indicators, the electrical load or reaction torque on the engine 112 may be calculated or estimated. Additionally, at least one sensing device 132 may be positioned adjacent to the shaft 142, for example, to sense or measure a performance indicator between the engine 112 and the first propulsion assembly 116, the shaft 142 mechanically coupling the engine 112 with the first propulsion assembly 116. The performance indicator may be an output torque of the engine 112 (or an input torque to the first propulsion assembly 116), a rotational speed of the shaft 142, or other suitable parameter. Like other performance indicators, performance indicators sensed or measured between the engine 112 and the first propulsion assembly 116 may be used to predict or estimate the electrical load on the engine 112.

The control system 130 also includes one or more controllers 134 configured to control various components of the hybrid electric propulsion system 110. The one or more controllers 134 are communicatively coupled with various components of the hybrid electric propulsion system 110. For this embodiment, one or more controllers 134 are communicatively coupled with one or more sensors 132, first electric machine 114, second electric machine 116, first power converter 118, second power converter 124, first propulsion assembly 116, second propulsion assembly 128, energy storage device 122, clutch 115, and engine 112, and more specifically with one or more engine controllers 136, the one or more engine controllers 136 configured to control engine 112. The engine controller 136 may be, for example, an Electronic Engine Controller (EEC) or an Electronic Control Unit (ECU) equipped with Full Authority Digital Engine Control (FADEC). The engine controller 136 includes various components for performing various operations and functions, such as for controlling various variable geometry components and controlling fuel flow to the combustor.

The one or more controllers 134 and the one or more engine controllers 136 may each include one or more processors and one or more memory devices. The one or more processors may include any suitable processing device, such as a microprocessor, microcontroller, integrated circuit, logic device, and/or other suitable processing device. The one or more memory devices may include one or more computer-readable media, including but not limited to non-transitory computer-readable media, RAM, ROM, hard drives, flash drives, and/or other memory devices.

The one or more memory devices may store information accessible by the one or more processors, including computer-readable instructions that may be executed by the one or more processors. The instructions may be any set of instructions that, when executed by one or more processors, cause the one or more processors to perform operations. The instructions may be software written in any suitable programming language, or may be implemented in hardware. Additionally, and/or alternatively, the instructions may be executed in logically and/or virtually separate threads on the processor.

The memory device may further store data that may be accessed by the one or more processors. For example, the data may include sensor data collected from various sensing devices 132 of the hybrid electric propulsion system 110. In particular, the data may include one or more performance indicators indicative of the electrical load on the first motor 114. The data may also include other data sets, parameters, outputs, information, etc. shown and/or described herein.

One or more controllers 134 and engine controllers 136 may each include a communication interface for communicating with each other, for example, with other components of aircraft 100. The communication interface may include any suitable components for interfacing with one or more networks, including, for example, a transmitter, a receiver, a port, a controller, an antenna, network interface components, and/or other suitable components. The one or more controllers 134 and the one or more engine controllers 136 may be communicatively coupled with a communication network of the aircraft 100. The communication network may include, for example, a Local Area Network (LAN), a Wide Area Network (WAN), a SATCOM network, a VHF network, an HF network, a Wi-Fi network, a WiMAX network, a gatelink network, and/or any other suitable communication network for transmitting messages to aircraft 100 and/or from aircraft 100, such as to a cloud computing environment and/or off-board computing systems. Such a network environment may use a variety of communication protocols. The communication network may include a data bus or a combination of wired and/or wireless communication links. The communication network may also be coupled to one or more controllers 134,136 by one or more communication cables or by wireless means.

As described above, the one or more controllers 134 are configured to control various components of the hybrid electric propulsion system 110 to take corrective action when rapid electrical load changes occur on the engine 112. The rapid electrical load change may be a load decrease or a load increase. In the event of a load reduction, one or more electrical components of the hybrid electric propulsion system 110 may fail or otherwise be controlled down-line. For example, the first motor 114 may fail, the second motor 126 may fail, etc. When this occurs, the electrical load on the engine 112 is rapidly reduced, or in other words, the torque resisting rotation of the output shaft of the engine 112 is rapidly reduced. This rapid change may cause a number of problems as previously described, such as overspeed, thrust asymmetry, etc. of the first motor 114. In the event of an increase in load, the power required by the power consumers or loads electrically coupled with the first electric machine 114 cannot be delivered by the first electric machine 114. That is, the required power is greater than the available power that can be generated by the first electric machine 114. The following provides an exemplary manner in which the control system 130 may take corrective action in the event of a rapid electrical load change.

In some embodiments, the one or more controllers 134 are configured to receive one or more performance indicators from the one or more sensing devices 132 indicative of the electrical load on the engine 112. For example, a sensing device 132 located adjacent to 138 that mechanically couples the engine 112 and the first electric machine 114 may sense a performance indicator indicative of the torque input to the first electric machine 114. Another sensing device 132 positioned adjacent to the shaft 140 mechanically coupling the second electric machine 126 with the second propulsion assembly 128 may sense a performance indicator indicative of the torque output of the second electric machine 126. Based on these performance indicators, and possibly other aspects, one or more controllers 134 determine the electrical load on the engine 112.

Next, the one or more controllers 134 are configured to determine whether there is a change in electrical load based at least in part on the one or more performance indicators. In some embodiments, the one or more controllers 134 are configured to determine an electrical load rate indicative of an electrical load of the engine 112 over a predetermined time interval based at least in part on the one or more performance indicators. Thereafter, the one or more controllers 134 are configured to compare the electrical load rate to a predetermined rate threshold. In such embodiments, the one or more controllers 134 determine that there is an electrical load change if the electrical load rate exceeds a predetermined rate threshold. In other embodiments, one or more controllers 134 are configured to receive status information from controllers 118 and/or 124 indicating a fault.

FIG. 3 provides a plot of torque load on an engine as a function of time, and depicts one example manner in which a determination of whether there is a change in electrical load may be based, at least in part, on one or more performance indicators. As shown, for a first predetermined time interval Int.1 from time T1 to time T2, the one or more controllers 134 determine an electrical load rate indicative of an electrical load on the engine 112 based at least in part on the one or more performance indicators. First predetermined time interval int.1 may be a single time step or multiple time steps, e.g., on the order of milliseconds or microseconds, for one or more controllers 134. In this example, various torque inputs are provided to the controller 134, and based on the torque inputs, the controller 134 calculates the total electrical load on the engine 112. During the first predetermined time interval int.1, the electrical load on the engine 112 remains constant, and thus, there is no change in the electrical load on the engine 112. Therefore, the control system 130 need not take corrective action during the first predetermined time interval int.1.

For a second predetermined time interval int.2 from time T2 to time T3, the one or more controllers 134 again determine an electrical load rate indicative of the electrical load on the engine 112 based at least in part on the one or more performance indicators. Second predetermined time interval int.2 may span the same amount of time as first predetermined time interval int.1. Based on the torque input, the controller 134 calculates the total electrical load on the engine 112. During the second predetermined time interval int.2, the electrical load on the engine 112 is reduced, and therefore, the load on the engine 112 is reduced. Accordingly, the one or more controllers 134 compare the electrical load rate to the predetermined rate threshold RT. For this embodiment, the predetermined rate threshold RT is set such that if the electrical load rate does not exceed the threshold, the reduction in electrical load on the engine does not cause overspeed damage to the first electrical machine 114 and any resulting thrust asymmetry does not result in a significant hazard to the aircraft 100. As shown in fig. 3, the torque load on the engine 112 decreases beyond a predetermined rate threshold RT (the rate of torque load on the engine 112 decreases at a faster rate than the predetermined rate threshold RT allows). Thus, in this example, the one or more controllers 134 determine that there is an electrical load change when the electrical load rate exceeds the predetermined rate threshold RT. It should be understood that other methods may be used to determine that there is an electrical load change. For example, the determination may be made based on other performance parameters such as bus power, engine/motor/propeller speed, or status indications provided by the controller 118 or 124.

Thereafter, the one or more controllers 134 are configured to generate a control action in response to the electrical load change if there is a rapid electrical load change. As will be explained further below, the appropriate control action response to an electrical load change on the engine 112 depends on whether the electrical load change is a load decrease (e.g., as shown in fig. 3) or a load increase.

Referring again to fig. 2, in the event of a load reduction or load drop, the one or more controllers 134 may generate a plurality of different control actions in accordance with exemplary aspects of the present disclosure. As one example, as shown in FIG. 2, the control system 130 includes a fuel control device 150, the fuel control device 150 configured to selectively control a fuel flow FF to the engine 112. For example, the fuel control device 150 may be positioned along a fuel line between a burner assembly of the engine 112 and a fuel tank 152. The fuel control device 150 is communicatively coupled with one or more controllers 134. The fuel control device 150 may be movable between a closed position and an open position based on one or more control signals from one or more controllers 134. The fuel control device 150 may be movable between an infinite number of open and closed positions, such as through the use of proportional control valves, or may be switchable between a single open and closed position. If the electrical load change is a decrease in load on the engine 112, the one or more controllers 134 are configured to activate the fuel control device 150 to decrease the fuel flow FF to the engine 112 when generating the control action. For example, the one or more controllers 134 may activate the fuel control device 150 to move to a closed position based on one or more control signals, e.g., to reduce the fuel flow FF to the engine 112. By reducing the fuel flow to the engine 112, the output torque of the engine 112 is reduced, allowing the engine output torque to drop to more quickly match the torque load (or electrical load) on the engine 112.

In some embodiments, additionally or alternatively to reducing the fuel flow FF to the engine 112 via the fuel control device 150, the control system 130 includes an electric brake system 160. For the embodiment shown in fig. 2, electric brake system 160 includes a load or resistor bank 162 comprised of a plurality of resistors and/or other dissipative elements. Resistor bank 162 may be positioned within fuselage 106 or at some other suitable location. Electric brake system 160 also includes an electric switching device 164 for selectively electrically coupling first electric machine 114 (or the DC bus) and resistor bank 162. For this embodiment, the electrical switching device 164 is located within the first power converter 118. In this manner, the first power converter 118 may direct all, some, or some of the power generated by the first motor 114 to the resistor bank 162.

In such embodiments, if the electrical load change is a decrease in load on the engine 112, the one or more controllers are configured to activate the electrical switching device 164 to electrically couple the first electric machine 114 and the resistor bank 162 during the transient period to direct electrical power from the first electric machine 114 to the resistor bank 162 when the control action is generated. In this manner, the resistive electrical load on the engine 112 may be quickly and automatically increased to accommodate the loss or drop in generator electrical load on the engine 112. Notably, the electrical switching apparatus 164 may be switched almost instantaneously (e.g., within a few microseconds) when the controller 134 determines that there is a reduction in electrical load. Thus, excess power generated by the first electric machine 114 may be dissipated via the resistor bank 162. This may prevent overspeeding of first motor 114 and reduce thrust asymmetry, for example. Although the output torque of the second electric machine 126 may be reduced to some extent, the overall thrust asymmetry is reduced by momentarily sinking the power from the engine 112. In particular, electrical power may be directed to the resistor bank 162 for a transient period of time, which may be a time sufficient for the engine 112 to reduce its torque output to match the torque load of the electrical system on the engine 112. For example, the transient time period may be several seconds.

In other embodiments, additionally or alternatively to reducing the fuel flow FF to the engine 112 via the fuel control device 150, the control system 130 includes a physical braking system 170 mechanically coupled with the engine 112. For the embodiment shown in fig. 2, the physical braking system 170 includes a disc brake 172 mechanically coupled to the engine 112. Although the disc brake 172 is shown as being located rearward or aft of the engine 112 in fig. 2, the disc brake 172 may be located in other suitable locations as long as the disc brake 172 is mechanically coupled with the engine 112. For example, the disc brake 172 may be positioned at a front portion of the engine 112, e.g., between the engine 112 and the first electric machine 114. In such embodiments, if the rapid electrical load change is a reduction in load on the engine 112, the one or more controllers are configured to activate the physical braking system 170 to reduce the torque output of the engine 112 during the transient period when the control action is generated. By placing a load on the engine 112 via the physical braking system 170, the engine 112 output torque is reduced, allowing the engine output torque to drop to more quickly match the torque load (or electrical load) on the engine 112. Additionally, the physical braking system 170 may prevent overspeeding of the engine 112 and/or the first electric machine 114.

In some further embodiments, additionally or alternatively to reducing the fuel flow FF to the engine 112 via the fuel control device 150, the control system 130 may utilize the energy storage device 122 electrically coupled with the first electric machine 114 to account for rapid electrical load changes. In some embodiments, if the rapid electrical load change is a reduction in load on the engine, the energy storage device 122 is configured to receive an amount of excess electrical power from the first electrical machine 114 during the transient period. The transient time period may be a time sufficient for the engine 112 to reduce its torque output to match the torque load of the electrical system on the engine 112. The first power converter 118 may include a switching element that directs all, some, or some of the power generated by the first electric machine 114 to the energy storage device 122. For example, some power may be directed to electric brake system 170, and some power may be directed to energy storage device 122. In other embodiments, switching elements may be positioned along power bus 120 for selectively directing excess power to energy storage device 122.

In further embodiments, additionally or alternatively to reducing the fuel flow FF to the engine 112 via the fuel control device 150, the control system 130 may utilize one or more variable geometry components 180 of the engine 112 to account for rapid electrical load changes, for example, by reducing the torque output of the engine 112. The one or more variable geometry components 180 of the engine 112 may include variable guide vanes, such as variable inlet guide vanes and variable outlet guide vanes, as well as other variable airfoil surfaces that vary or affect the mass flow of the gas path through the engine. Other variable geometry components may include exhaust valves, one or more High Pressure Turbine Active Clearance Control (HPTACC) valves, Low Pressure Turbine Active Clearance Control (LPTACC) valves, Core Compartment Cooling (CCC) valves, booster anti-icing (BAI) valves, nacelle anti-icing (NAI) valves, Start Bleed Valves (SBV) (Start bleed valves), Transient Bleed Valves (TBV), Modulated Turbine Cooling (MTC) valves, and/or combination valves. In such embodiments, if the rapid electrical load change is a decrease in load on the engine 112, the one or more controllers 134 are configured to activate the one or more variable geometry components 180 of the engine 112 such that the engine 112 operates in a less efficient manner during transient periods.

For example, upon determining that a rapid load reduction has occurred, the one or more controllers 134 may communicate with the one or more engine controllers 136, e.g., to instruct the one or more engine controllers 136 to actuate the one or more variable geometry components 180 of the engine 112 in a less efficient manner. By operating the engine 112 in a less efficient manner, the torque output of the engine 112 will decrease, and therefore, the torque output of the engine 112 will decrease more quickly so that it more quickly matches the torque load on the engine 112. The transient time period may be a time sufficient for the engine 112 to reduce its torque output to match the torque load of the electrical system on the engine 112. When the torque output of the engine 112 matches the torque load of the electrical system on the engine 112, the one or more controllers 134 may communicate with the engine controller 136 to control the one or more variable geometry components 180 to operate in a more efficient manner, for example, to increase the efficiency of the engine 112.

In other embodiments, additionally or alternatively to reducing the fuel flow FF to the engine 112 via the fuel control device 150, the control system 130 may utilize one or more of the propulsion assemblies 116,128 to account for rapid electrical load changes, for example, by reducing the torque load on the engine 112. In such embodiments, if the rapid electrical load change is a reduction in load on the engine 112, the one or more controllers 134 are configured to control the first propulsion assembly 116, the second propulsion assembly 128, some other propulsion assembly mechanically coupled with the electrical machine of the electrical system, or some combination thereof, in a less efficient manner during the transient period. For example, in some embodiments, first and second propulsion assemblies 116,128 are variable pitch propeller assemblies, each propeller assembly having a plurality of blades that are adjustable in unison by a plurality of pitches or blade angles, such as by activation of an actuation mechanism. Pitching of the blades may cause the propeller assembly to produce more or less thrust. In the event of a reduced load on the engine 112, the controller 134 may communicate directly or indirectly with the activation mechanism of the propulsion assembly 116,128 to pitch the blades to a more rough or flat angle. In this manner, the torque load on the engine 112 may be increased, and thus some excess torque output from the engine 112 may be accounted for by the propulsion assembly 116,128. Advantageously, adjusting one or more propulsion assemblies may prevent the engine/motor from overspeeding.

In some embodiments, additionally or alternatively to reducing the fuel flow FF to the engine 112 via the fuel control device 150, the control system 130 may reduce the efficiency of the first electric machine 114, the second electric machine 126, or some other electrical consumer or load electrically coupled with the first electric machine 114 to account for the drop in electrical load on the engine 112. In such embodiments, if the rapid electrical load change is a reduction in load on the engine 112, the one or more controllers 134 are configured to control at least one of the first electric machine 114, the second electric machine 126, other electrical power consuming devices (e.g., power converters 118,124, other electric machines, etc.) electrically coupled with the first electric machine 114, or some combination thereof, to operate in a less efficient manner during the transient period. In this manner, the transient excess power generated by the first electric machine 114 may be dissipated as heat via one or more electric machines or power consumers electrically coupled with the first electric machine 114. As an example, the phase or current advance angle, the magnitude of the phase current, or some other parameter known to affect the efficiency of one or more electric machines may be controlled to operate the one or more electric machines in a reduced efficiency mode during an excess power transient. The transient time period may be a time sufficient for the engine 112 to reduce its torque output to match the torque load of the electrical system on the engine 112. By operating the first electric machine 114 in a less efficient or high power manner, the second electric machine 16, or some other electrical consumer or load, an electrical load may be placed on the engine 112 to account for the load reduction. Such transient electrical loads may be placed on the engine 112 almost instantaneously (e.g., within microseconds of determining that there is a reduction in load on the engine 112).

In some embodiments, the first electric machine 114 is a generator (e.g., 214 of fig. 4) and the second electric machine 126 is an electric motor (e.g., 206 of fig. 4) electrically coupled to the generator 214. If the rapid electrical load change is a reduction in load on the engine 112 (or 202 of FIG. 4), the one or more controllers 134 (or 220 of FIG. 4) are also configured to control at least one of the electric motor 206 or other electrical consuming load (which may also be an electric motor) to operate in a less efficient or high power manner.

In some embodiments, additionally or alternatively to reducing the fuel flow FF to the engine 112 via the fuel control device 150, the control system 130 may direct transient excess electrical power to one or more accessory loads 190 (fig. 1) of the aircraft 100. Exemplary accessory loads 190 may include an air conditioning unit, a pump or fan, a display, a data processing unit, a communication unit, other subsystems of the aircraft 100, some combination thereof, and the like. The control system 130 of the hybrid electric propulsion system 110 includes an electric switching device 192 for selectively electrically coupling the first electric machine 114 and one or more accessory loads 190 of the aircraft 100. For this embodiment, the electrical switching device 192 is located within the second power converter 124. In this manner, the second power converter 124 may direct all, a portion, or some of the power generated by the first motor 114 to the one or more accessory loads 190, e.g., as AC current. In such an embodiment, if the rapid electrical load change is a decrease in load on the engine 112, the one or more controllers 134 are configured to activate the electrical switching device 192 to electrically couple the first electric machine 114 and the one or more accessory loads 190 during the transient period to direct electrical power from the first electric machine 114 to the one or more accessory loads 190. By increasing the electrical load on the engine 112 during transient periods using the accessories 190 of the aircraft 100, the engine 112 may more quickly match its torque output to the torque load of the electrical system on the engine 112. It will be appreciated that the electrical switching apparatus 192 may be located in other suitable locations of the electrical system, and that the power switching and steering may be accomplished in any suitable manner and in any suitable physical location.

In the event of a load increase, the one or more controllers 134 may generate a plurality of different control actions in accordance with exemplary aspects of the present disclosure. Whether there is a load increase may be determined by the one or more controllers 134, for example, by determining an electrical load rate indicative of an electrical load on the engine 112 over a predetermined time interval based at least in part on one or more performance indicators, and then comparing the electrical load rate to a predetermined rate threshold. In such embodiments, if the electrical load rate exceeds the predetermined rate threshold (e.g., by having a slope greater than the slope of the predetermined threshold slope), the one or more controllers 134 determine that there is an electrical load change, and in particular, that there is a load increase. It should be appreciated that other suitable methods may be used to determine that there is a change in electrical load. For example, the determination may be made based on other performance parameters such as bus power, engine/motor/propeller speed, or status indications provided by the controller 118 or 124.

In some embodiments, if the rapid electrical load change is an increase in load on the engine 112 and the increase in load is caused by a failure of the first electric machine 114, the one or more controllers 134 are configured to control delivery of electrical power from the energy storage device 122 to the second electric machine 126 during the transient period. In this manner, power may be delivered to the second motor 126 despite a failure of the first motor 114 such that the second motor 126 may drive the second propulsion assembly 128. Thus, thrust asymmetry is reduced or eliminated.

In other embodiments, if the rapid electrical load change is an increase in load on the engine 112 and an intentionally demanded load increase, the one or more controllers 134 are also configured to control delivery of electrical power from the energy storage device 122 to the second electrical machine 126 during the transient period. For example, if the power required and consumed by the second electric machine 126 to drive the second propulsion assembly 128 spikes (e.g., during vertical takeoff or landing, hover maneuvers, etc.) such that the electrical load on the engine 112 increases rapidly, the one or more controllers 134 control the energy storage device 122 to deliver power to the second electric machine 126 for a transient period of time to achieve the desired power demand. In this manner, the control system 130 may facilitate the electrical power required to meet the electrical system of the hybrid electric propulsion system 110.

Further, in some embodiments, the first motor 114 may be mechanically coupled with the first propulsion assembly 116 (e.g., as shown in fig. 2). The first motor 114 may be directly mechanically coupled with the first propulsion assembly 116, or an additional clutch (not shown) may be disposed therebetween to decouple the first motor 114 from the first propulsion assembly 116. In the event of a catastrophic failure of the engine 112, the clutch 115 may be controlled to disengage the first electric machine 114 from the engine 112, and the first electric machine 114 may receive electrical power from the energy storage device 122. A first motor 114, acting as a motor, may drive a first propulsion assembly 116. Thus, in some embodiments, the hybrid electric propulsion system 110 may switch to an all-electric system.

Further, in some cases, the first motor 114 may fail. In the event of a failure of the first electric machine 114, the amount of electrical power entering the second electric machine 126 decreases, and the torque load on the engine 112 also decreases. In some embodiments, the lack of power to the second electric machine 126 may be compensated for by drawing power from the energy storage device 122. Other methods already mentioned may be used to prevent the engine 112 and the first propulsion assembly 116 from overspeeding.

FIG. 4 provides a schematic illustration of another exemplary hybrid electric propulsion system 200 for an aircraft, according to an exemplary embodiment of the present disclosure. The hybrid electric propulsion system 200 of fig. 4 is configured for driving propulsion and supplying power to any suitable aircraft, such as unmanned aerial vehicles configured for vertical takeoff and landing and hover maneuvers. The exemplary hybrid electric propulsion system 200 of fig. 4 is configured in a similar manner as the hybrid electric propulsion system 110 of fig. 1 and 2, except as described below.

In contrast to the hybrid electric propulsion system 110 of fig. 1 and 2, the hybrid electric propulsion system 200 of fig. 4 includes a torque source or engine 202 that is not mechanically coupled with the propulsion assembly. Rather, for this embodiment, the engine 202 is configured to drive one or more generators 204. More specifically, the engine 202 is configured to drive a pair of generators 204, each generator 204 being mechanically coupled to the engine 202. Similar to the first electric machine 114 of fig. 1 and 2, the generator 204 is configured to generate electrical power when driven by the engine 202. In some alternative embodiments, the hybrid electric propulsion system 200 may include one generator or more than two (2) generators.

As further shown in FIG. 4, a plurality of electrical power consuming devices 206 are electrically coupled with the generator 204 via an electrical bus 210. For this embodiment, the electrical power consuming device 206 includes a plurality of electric motors configured to receive electrical power from one of the generators 204 and generate an output torque to drive a propulsion assembly 208 mechanically coupled thereto. For example, the propulsion assembly 208 may be a fan, rotor, or other suitable propulsion device. Further, for this embodiment, some electrical consumers 206 are powered by the generator 204, but are not mechanically coupled with the propulsion assembly 208. For example, such an electrical consumer 206 may power a display, a controller 220 of the hybrid electric propulsion system 200, an avionics system, and the like. The energy storage device 214 is also electrically coupled with, e.g., receives electrical energy from and delivers electrical energy to, the generator 204 and the electrical consumer 206. Although not shown, various power converters may be included in the hybrid-electric powertrain 200, for example, to convert the generated electrical power into a desired form.

A plurality of sensing devices 212 are positioned to detect one or more performance indicators indicative of an electrical or torque load on the engine 202. As one example, the one or more sensing devices 212 may be a speed sensor configured to sense the rotational speed of a shaft or rotating component located in proximity thereto. As another example, the one or more sensing devices 212 may be torque sensors configured to measure the torque of a shaft or rotating component in the vicinity thereof. As yet another example, the one or more sensing devices 212 may be configured to measure or sense an electrical parameter indicative of a torque load on the engine 112, such as current flowing through an electric machine, voltage across the machine, or some other parameter. These performance indicators may be directed to controller 220, and controller 220 may be multiple controllers or a single controller, such that the torque load or electrical load on engine 202 may be calculated or estimated. According to an exemplary aspect disclosed herein, the hybrid electric propulsion system 200 includes a control system 216, the control system 216 configured to take rapid and automatic corrective action in the event of rapid electrical load changes on the engine 202, e.g., similar to that described above with respect to the control system 130 of the hybrid electric-based system 110 of fig. 1 and 2. For example, the controller 220 of the control system 216 may control various components of the hybrid electric propulsion system 200 in the same or similar manner as described above. Some or all of the components described above with respect to the control system 130 of the hybrid-electric propulsion system 110 of fig. 1 and 2 may be incorporated into the control system 216 of the hybrid-electric propulsion system 200.

FIG. 5 provides a flowchart of an exemplary method (300) for operating a hybrid electric propulsion system for an aircraft. For example, the hybrid electric propulsion system and the aircraft may be the hybrid electric propulsion system 110, and the aircraft may be the aircraft 100 of fig. 1 and 2. The method (300) is also applicable to the hybrid electric propulsion system 200 of fig. 4. In this context, reference numerals will be utilized below to describe the hybrid electric propulsion system 110 and the aircraft 100 described above and various features thereof.

At (302), the method (300) includes receiving, from one or more sensing devices located on the aircraft, one or more performance indicators indicative of an electrical load on an engine, the engine mechanically coupled with an electric machine configured to generate electrical power when driven by the engine, the electric machine electrically coupled with one or more electrical power consuming devices, wherein at least one of the one or more electrical power consuming devices is mechanically coupled with the propulsion assembly. For example, the sensing device 132 positioned on the aircraft 100 may sense one or more performance indicators indicative of the electrical load on the engines 112. The performance indicator may be a torque or speed of a rotating component, an electrical parameter (e.g., current, voltage, etc.) of an electric machine or electrical consumer, or some other indicator of an electrical load or torque load on the engine 112. Some of the signals or measurements may then be directed to one or more controllers 134 for processing. In this manner, the electrical load on the engine 112 may be determined. The combination of the measurements and estimates may be used to calculate the electrical load on the engine 112.

At (304), the method (300) includes determining whether there is a reduction in electrical load on the engine based at least in part on the one or more performance indicators. For example, in some embodiments, the one or more controllers 134 are configured to determine an electrical load rate indicative of an electrical load on the engine 112 over a predetermined time interval based at least in part on the one or more performance indicators. Thereafter, the one or more controllers 134 are configured to compare the electrical load rate to a predetermined rate threshold. In such embodiments, the one or more controllers 134 determine that there is an electrical load change if the electrical load rate exceeds a predetermined rate threshold. Fig. 3 and the accompanying text illustrate an exemplary manner in which it may be determined whether there is a rapid electrical load reduction. In some further embodiments, the torque demanded by the electrical system on the engine 112 may also be considered. Thus, in addition to checking whether there is a reduction in electrical load based on comparing the electrical load on the engine 112 to a predetermined rate threshold, the required torque of the system may be used to determine whether there is a reduction in electrical load.

At (306), the method (300) comprises: generating a control action in response to the electrical load reduction if there is an electrical load reduction on the engine, wherein generating the control action comprises reducing fuel flow to the engine. For example, as previously described, the control system 130 of the hybrid electric propulsion system 110 may include a fuel cutoff device or fuel control device 150 configured to selectively control the fuel flow FF to the engine 112 (e.g., as shown in fig. 2). Upon determining that there is a load reduction, the controller 134 communicatively coupled with the fuel control device 150 activates the fuel control device 150 to move toward the closed position to shut off the flow of fuel, FF, to the engine 112. By reducing the fuel flow to the engine 112, the engine 112 output torque is reduced, allowing the engine output torque to drop faster to match the torque load (or electrical load) on the engine 112.

In some embodiments of the method (300), the controlling act further includes delivering an amount of excess electrical power from the electric machine to the electric brake system during the transient period. For example, in some embodiments, control system 130 includes electric brake system 160 of fig. 2, which includes a resistor bank 162 comprised of a plurality of resistors and/or other dissipative elements. Electric brake system 160 also includes an electric switching device 164 for selectively electrically coupling first electric machine 114 and resistor bank 162. Upon determining that there is a load reduction, controller 134, which is communicatively coupled with electrical switching device 164, activates electrical switching device 164 to electrically couple the electric machine and resistor bank 162 to direct power from the electric machine to resistor bank 162. In this manner, the electrical load on the engine 112 may be quickly and automatically increased to accommodate the loss or drop in electrical load on the engine 112. Thus, excess power generated by the motor may be dissipated via the resistor bank 162. The transient period of time during which the resistor bank 162 receives power may be the time during which the engine 112 decelerates to match its output torque to the electrical load on the engine 112.

In some embodiments of the method (300), the controlling act further includes delivering an amount of excess power from the electric machine to one or more accessory loads of the aircraft during the transient period. For example, in some embodiments as shown in fig. 2, one or more accessory loads 190 are electrically coupled with an electric machine, which may be, for example, the first electric machine 114 of the hybrid electric propulsion system 110 of fig. 2. The accessory load may be any electrical consuming load of aircraft 100 (FIG. 1). Exemplary accessory loads 190 may include an air conditioning unit, a pump or fan, a display, a data processing unit, a communication unit, other subsystems of the aircraft 100, some combination thereof, and the like. In some embodiments, the power management device is electrically coupled with the electric machine and is configured to selectively distribute excess load to accessory loads 190 of the aircraft 100. In the event that a load reduction is determined to exist, the controller 134 communicatively coupled with the power management device may instruct the power management device to distribute excess power to one or more accessory loads 190 of the aircraft 100. In this manner, the electrical load on the engine 112 may be quickly and automatically increased to accommodate the loss or drop in electrical load on the engine 112. The transient period of time during which the accessory load 190 receives power may be a period of time during which the engine 112 is decelerating to match its output torque to the electrical load on the engine 112.

In some embodiments of the method (300), the controlling act further includes controlling one or more variable geometry components of the engine to move such that the engine operates in a less efficient manner during the transient period. For example, in some embodiments as shown in FIG. 2, the engine 112 may include one or more variable geometry components 180, such as variable guide vanes positioned along one or more gas paths of the engine 112. In such embodiments, upon determining at (304) that a rapid load reduction on the engine 112 has occurred, based on one or more command signals from the controller 134, the one or more engine controllers 136 control the one or more variable geometry components 180 of the engine 112 to actuate or move such that the engine 112 operates in a less efficient manner during transient periods. For example, inlet guide vanes of a compressor of a gas turbine engine may be moved such that the vanes impart less efficient rotational motion to the airflow passing through the vanes, thereby reducing the torque output of the engine more quickly. As previously described, by operating the engine 112 in a less efficient manner, the torque output of the engine 112 will decrease, and therefore, the torque output of the engine 112 will decrease more quickly, such that it more quickly matches the torque load on the engine 112. The transient time period may be a time sufficient for the engine 112 to reduce its torque output to match the torque load of the electrical system on the engine 112.

In some embodiments of the method (300), the controlling act further includes controlling the electric machine and/or at least one of the one or more electrical consumers or loads to operate at a less efficient or high power during the transient time period. For example, the motor may be the first motor 114 of fig. 1 and 2, and the electrical consumers may include the second motor 126 and other loads. In the event that a load reduction is determined to exist, the one or more controllers 134 control at least one of the first electric machine 114, the second electric machine 126, other electrical power consuming devices (e.g., power converters 118,124, other electric machines, etc.) electrically coupled with the first electric machine 114, or some combination thereof, to operate at a less efficient or high power during transient periods. In this manner, excess electrical power generated by the first electrical machine 114 during the transient state may be dissipated as heat via one or more electrical machines or electrical power consuming devices electrically coupled with the first electrical machine 114. The transient time period may be a time sufficient for the engine 112 to reduce its torque output to match the torque load of the electrical system on the engine 112.

In some further embodiments of the method (300), the controlling act further includes controlling the propulsion assembly in a less efficient manner during the transient time period. For example, the propulsion assembly may be the first propulsion assembly 116 of fig. 1 and 2. In the event that it is determined that there is a reduction in load, the one or more controllers 134 may control the first propulsion assembly to operate at a lower efficiency, for example by pitching the blades to a coarser or flat angle (i.e., an angle as opposed to a feathering angle). In this manner, the torque load on the engine 112 may be increased, and thus some excess torque output from the engine 112 may be accounted for by the propulsion assembly 116. Of course, more than one propulsion assembly may be operated at a time with less efficiency, such as in the case of the propulsion assembly 208 of FIG. 4. Advantageously, adjusting one or more propulsion assemblies may prevent the engine/motor from overspeeding.

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

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 languages of the claims.

Further aspects of the invention are provided by the subject matter of the following clauses:

1. a hybrid electric propulsion system for an aircraft, the hybrid electric propulsion system comprising: an engine; a first electric machine mechanically coupled with the engine and configured to generate electrical power when driven by the engine; one or more electrical power consuming loads electrically coupled with the first electrical machine and configured to receive electrical power from the first electrical machine, the one or more electrical power consuming loads including a second electrical machine; a propulsion assembly mechanically coupled with the second motor and configured to generate thrust when driven by the second motor; one or more sensing devices located on the aircraft for detecting one or more performance indicators indicative of electrical loads on the engines; one or more controllers communicatively coupled with the one or more sensors, the one or more controllers configured to: receiving the one or more performance indicators indicative of the electrical load on the engine from the one or more sensing devices; determining whether there is a change in electrical load based at least in part on the one or more performance indicators; and if there is a change in the electrical load, generating a control action in response to the change in the electrical load.

2. The hybrid electric propulsion system of any preceding claim, further comprising: a fuel control device configured to selectively control fuel flow to the engine, the fuel control device communicatively coupled with the one or more controllers, and wherein if the electrical load change is a decrease in load on the engine, the one or more controllers, in generating the control action, are configured to: activating the fuel control device to reduce the fuel flow to the engine.

3. The hybrid electric propulsion system of any preceding claim, further comprising: an electric brake system including a resistor bank and an electric switching arrangement for selectively electrically coupling the first electric machine and the resistor bank, and wherein if the electrical load change is a decrease in load on the engine, the one or more controllers, in generating the control action, are configured to: activating the electrical switching apparatus to electrically couple the first electric machine and the resistor bank to direct electrical power from the first electric machine to the resistor bank.

4. The hybrid electric propulsion system of any preceding claim, further comprising: a physical braking system mechanically coupled with the engine, and wherein if the electrical load change is a decrease in load on the engine, the one or more controllers, in generating the control action, are configured to: activating the physical braking system to reduce the torque output of the engine.

5. The hybrid electric propulsion system of any preceding claim, further comprising: an energy storage device electrically coupled with the first electric machine and electrically coupled with the second electric machine, and wherein the energy storage device is configured to receive an amount of excess power from the first electric machine if the electrical load change is a decrease in load on the engine.

6. The hybrid electric propulsion system of any preceding item, wherein if the electrical load change is a decrease in load on the engine, the one or more controllers are further configured to: activating one or more variable geometry components of the engine to operate the engine in a less efficient manner.

7. The hybrid electric propulsion system of any preceding claim, wherein the first electric machine is a generator and the second electric machine is an electric motor electrically coupled with the generator, and wherein, if the electrical load change is a decrease in load on the engine, the one or more controllers are further configured to: controlling at least one of the electric motor and the one or more electrical consuming loads to operate at a less efficient or high power.

8. The hybrid electric propulsion system of any preceding item, wherein the aircraft includes one or more accessory loads, and wherein the hybrid electric propulsion system further comprises: an electrical switching arrangement for selectively electrically coupling the first electric machine and the one or more accessory loads, and wherein, if the electrical load change is a reduction in load on the engine, the one or more controllers are further configured to: activating the electrical switching apparatus to electrically couple the first electric machine and the one or more accessory loads to direct power from the first electric machine to the one or more accessory loads.

9. The hybrid electric propulsion system of any preceding item, wherein if the electrical load change is an increase in load on the engine, the one or more controllers are further configured to: controlling an energy storage device to deliver power to the second electric machine during transient periods.

10. A method for operating a hybrid electric propulsion system for an aircraft, the method comprising: receiving one or more performance indicators indicative of an electrical load on the engine from one or more sensing devices located on the aircraft, the engine mechanically coupled with an electric machine configured to generate electrical power when driven by the engine, the electric machine electrically coupled with one or more electrical power consuming devices, wherein at least one of the one or more electrical power consuming devices is mechanically coupled with a propulsion assembly; determining whether there is a reduction in electrical load on the engine based at least in part on the one or more performance indicators; and if there is a decrease in the electrical load on the engine, generating a control action in response to the decrease in the electrical load, wherein generating the control action comprises decreasing fuel flow to the engine.

11. The method of any of the preceding claims, wherein the controlling act further comprises: delivering an amount of excess electrical power from the electrical machine to an electric brake system.

12. The method of any of the preceding claims, wherein the controlling act further comprises: providing an amount of excess power from the electric machine to one or more accessory loads of the aircraft.

13. The method of any of the preceding claims, wherein the controlling act further comprises: controlling one or more variable geometry components of the engine to move such that the engine operates in a less efficient manner.

14. The method of any of the preceding claims, wherein the controlling act further comprises: controlling at least one of the motor or the one or more power consuming devices to operate in a less efficient manner.

15. The method of any of the preceding claims, wherein the controlling act further comprises: the propulsion assembly is controlled in a less efficient manner during transient periods.

16. A hybrid electric propulsion system for an aircraft, the hybrid electric propulsion system comprising: an engine; a first propulsion assembly mechanically coupled with the engine and configured to generate thrust when driven by the engine; a first electric machine mechanically coupled with the engine and configured to generate electrical power when driven by the engine; a second electric machine electrically coupled with the first electric machine and configured to receive power from the first electric machine; a second propulsion assembly mechanically coupled with the second motor and configured to generate thrust when driven by the second motor; one or more sensing devices located on the aircraft for detecting one or more performance indicators indicative of electrical loads on the engines; one or more controllers communicatively coupled with the one or more sensors, the one or more controllers configured to: receiving the one or more performance indicators indicative of the electrical load on the engine from the one or more sensing devices; determining whether there is a change in electrical load based at least in part on the one or more performance indicators; and if there is a change in the electrical load, generating a control action in response to the change in the electrical load.

17. The hybrid electric propulsion system of any preceding claim, further comprising: the energy storage device is electrically coupled with the second electric machine, and wherein, if the electrical load change is caused by a failure of the first electric machine, the one or more controllers are further configured to: controlling delivery of electrical power from the energy storage device to the second electrical machine to drive the second propulsion assembly.

18. The hybrid electric propulsion system of any preceding claim, further comprising: a fuel control device configured to selectively control fuel flow to the engine, the fuel control device communicatively coupled with the one or more controllers, and wherein if the electrical load change is a decrease in load on the engine, the one or more controllers, in generating the control action, are configured to: activating the fuel control device to reduce the fuel flow to the engine.

19. The hybrid electric propulsion system of any preceding claim, wherein at least one of the one or more sensing devices is positioned proximate an output shaft mechanically coupling the engine with the first electric machine and positioned adjacent an output shaft mechanically coupling the second electric machine and the second propulsion assembly.

20. The hybrid electric propulsion system of any preceding item, wherein to determine whether the electrical load change is present, the one or more controllers are configured to: determining an electrical load rate indicative of the electrical load on the engine over a predetermined time interval based at least in part on the one or more performance indicators; and comparing the electrical load rate to a predetermined rate threshold; wherein the one or more controllers determine that the electrical load change exists if the electrical load rate exceeds the predetermined rate threshold.