GooHYBRID control system across multiple operating modes

Abstract

A lever for adjusting an output of a hybrid electric powerplant of an aircraft includes a lever configured to move through a range of positions. Movement of the lever adjusts the output of the hybrid electric powerplant between at least two operating modes. In a first subset of positions within the range of positions, the hybrid electric powerplant is configured to operate an engine having a mechanical output, output a first electrical energy from a motor / generator driven by the mechanical output of the engine, and drive a propulsion mechanism with the mechanical output of the engine. In a second subset of positions within the range of positions, the hybrid electric powerplant is configured to operate an engine having a mechanical output, receive a second electrical energy from the motor / generator, drive a mechanical output from the motor / generator using the second electrical energy, and drive a propulsion mechanism with the mechanical output.

Description

[Technical field]

[0001] CROSS-REFERENCE TO RELATED PATENT APPLICATIONS This application claims the benefit of each of U.S. Provisional Patent Applications Nos. 63 / 280,589 and 63 / 280,560, each filed on November 17, 2021, the entire contents of each of which are incorporated herein by reference in their entirety.

[0002] This application relates to a GooHYBRID control system across multiple operating modes. [Background technology]

[0003] There are various types of aircraft propelled using different types of propulsion mechanisms, such as propellers, turbine or jet engines, rockets, or ramjets. The different types of propulsion mechanisms may be driven in different ways. For example, some propulsion mechanisms, such as propellers, may be driven by internal combustion engines or electric motors. As such, the combination of the propulsion mechanisms and the methods of powering them are often designed specifically for a particular aircraft, so that the propulsion mechanisms and the methods of powering them meet the specifications necessary to properly and safely propel the aircraft. Summary of the Invention

[0004] In one embodiment, a control system for regulating an output of a hybrid electric powerplant of an aircraft includes an input of a controller configured to receive commands. The controller is configured to set an operating mode of the hybrid system based on the command. The operating mode includes an output mode of the hybrid electric powerplant. There are at least two operating modes. With a first command provided to the input, the hybrid electric powerplant is configured to operate an engine having a mechanical output, output a first electrical energy from a motor / generator driven by the mechanical output of the engine, and drive a propulsion mechanism with the mechanical output of the engine. With a second command received, the hybrid electric powerplant is configured to operate the engine having a mechanical output, receive a second electrical energy at the motor / generator, drive the mechanical output with the second electrical energy at the motor / generator, and drive the propulsion mechanism with the mechanical output.

[0005] In one embodiment, a lever for adjusting an output of a hybrid electric powerplant of an aircraft includes a lever configured to move through a range of positions. Movement of the lever adjusts the output of the hybrid electric powerplant between at least two operating modes. In a first subset of positions within the range of positions, the hybrid electric powerplant is configured to operate an engine having a mechanical output, output a first electrical energy from a motor / generator driven by the mechanical output of the engine, and drive a propulsion mechanism with the mechanical output of the engine. In a second subset of positions within the range of positions, the hybrid electric powerplant is configured to operate an engine having a mechanical output, receive a second electrical energy from the motor / generator, drive a mechanical output from the motor / generator using the second electrical energy, and drive a propulsion mechanism with the mechanical output.

[0006] In one embodiment, a thrust control system for regulating an output of a hybrid electric powerplant of an aircraft includes an input of a controller configured to receive a command. The controller is configured to set an operating mode of the hybrid system based on the command. The operating mode includes an output mode of the hybrid electric powerplant. There are at least two operating modes. Upon receiving a first command at the input, the hybrid electric powerplant is configured to operate an engine having a mechanical output and output a first electrical energy from a motor / generator driven by the mechanical output of the engine, the first electrical energy being output to an aircraft electric propulsion motor and to an aircraft battery. Upon receiving a second command at the input, the hybrid electric powerplant is configured to output a second electrical energy from the motor / generator, the second electrical energy being output to the aircraft electric propulsion motor instead of to the aircraft battery. [Brief description of the drawings]

[0007] [Figure 1A] FIG. 1 illustrates an example of a flexible architecture for an aerospace hybrid system in accordance with an illustrative embodiment. [Figure 1B] FIG. 1 illustrates an additional example of a flexible architecture for an aerospace hybrid system in accordance with an illustrative embodiment. [Figure 2A] FIG. 1 is a block diagram illustrating a first aircraft control system for use in a flexible architecture for an aerospace hybrid system in accordance with an illustrative embodiment. [Figure 2B] FIG. 1 is a block diagram illustrating a second aircraft control system for use in a flexible architecture for an aerospace hybrid system in accordance with an illustrative embodiment. [Diagram 3] FIG. 1 illustrates an example of a first aircraft that can use the flexible architecture for an aerospace hybrid system in accordance with an illustrative embodiment. [Figure 4]FIG. 1 illustrates an example of a second aircraft that can use the flexible architecture for an aerospace hybrid system in accordance with an illustrative embodiment. [Diagram 5] FIG. 10 illustrates an example of a third aircraft that may use the flexible architecture for an aerospace hybrid system in accordance with an illustrative embodiment. [Figure 6] 1 is a flowchart illustrating a first example method of using a flexible architecture for an aerospace hybrid system during different flight phases of an aircraft including a main propulsion propeller in accordance with an illustrative embodiment. [Figure 7] 10 is a flowchart illustrating a second example method of using a flexible architecture for an aerospace hybrid system during different flight phases of an aircraft including a main propulsion propeller in accordance with an illustrative embodiment. [Figure 8] FIG. 1 illustrates an example of a flexible architecture for an aerospace hybrid system having a flywheel in accordance with an illustrative embodiment. [Figure 9] FIG. 1 illustrates a perspective view of an example flexible architecture for an aerospace hybrid system in accordance with an illustrative embodiment. [Figure 10] 10 is a top view illustrating an example of the flexible architecture of FIG. 9 in accordance with an exemplary embodiment. [Figure 11] FIG. 10 is a side view illustrating an example of the flexible architecture of FIG. 9 in accordance with an exemplary embodiment. [Figure 12] FIG. 1 illustrates a perspective view of another example of a flexible architecture for an aerospace hybrid system in accordance with an illustrative embodiment. [Figure 13] FIG. 1 illustrates an example downstream and upstream components for propelling an aircraft in accordance with an example embodiment. [Figure 14] FIG. 1 is a schematic diagram of an example system for providing a stable voltage to a direct current (DC) bus in accordance with an example embodiment. [Figure 15] 1 is a flowchart illustrating an example method for maintaining a stable DC bus voltage based on communications from an aircraft-level controller in accordance with an illustrative embodiment. [Figure 16] 1 is a flowchart illustrating an example method for maintaining a stable DC bus voltage based on measurements by a hybrid generator set-level controller in accordance with an illustrative embodiment. [Figure 17] FIG. 1 illustrates an example hybrid control system spanning multiple operating modes according to an example embodiment. [Figure 18] FIG. 1 illustrates an example operational mode in which an example hybrid architecture may be controlled in accordance with an example embodiment. [Figure 19] FIG. 1 is a schematic diagram of an example computing environment in accordance with an illustrative embodiment. DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

[0008] Various embodiments of thrust control across multiple operating modes of a hybrid generator set are described herein. Various types of operating modes of a hybrid generator set are described herein below, such as in the section entitled "Hybrid Generator Sets and Their Operating Modes." Various flight modes may be advantageous for use with an aircraft, but may be complicated for a human or computer / controller to operate. For example, a hybrid power plant system may include multiple modes, such as a parallel hybrid mode that combines a direct output shaft with a generator output, and it may be difficult or impossible for a pilot or on-board operator to efficiently use all the modes and switch between the modes. In other words, the pilot or other operator (including, for example, a computerized operator or an automated operator) may be focused on the thrust required to satisfy a mission or a particular phase of a mission, and it may be too mentally taxing for the pilot or other operator to think about power levels. In other words, an automated system or a human pilot / operator may prefer to provide specific instructions regarding overall thrust rather than providing or having input instructions specifying transitions between multiple operating modes.

[0009] Thus, described herein is a one-lever thrust lever design that can span at least two operating modes of a hybrid power plant. For example, in a first (low) operating range, a single lever movement can result in a blend of thrust outputs, including: (1) mechanical shaft power from the engine ranging from low or zero to most / full power depending on the position of the lever within the first (low) operating range, and (2) electrical power output by the motor / generator to the electric bus, where the generator is driven by the mechanical shaft power and thus the electrical power generated ranges from high or maximum power (e.g., when all of the mechanical shaft power is converted to electrical energy) to low or zero electrical output (e.g., less or no electrical power is generated and delivered on the electric bus) as the lever approaches the upper end of the first (low) operating range.

[0010] Once the lever has moved out of the first (low) operating range, it can be moved to a second (high) operating range. At this point, the generator can stop using a portion of the machine shaft power to generate electrical power and instead receive electrical power (e.g., from a battery) to further drive the machine shaft (e.g., power to the machine shaft greater than the engine could achieve alone). In other words, from this point on, as the lever is moved further, the thrust level automatically switches the motor / generator into motor mode, drawing electrical power from the DC bus to add to the shaft thrust provided to the propeller, fan, or gearbox. In such an embodiment, in the second (high) range, the engine output is at a constant high or maximum level, and the electrical power drawn from the DC bus can range from zero or a low level at the lower limit of the second (high) range of the lever to a high or maximum level at the upper limit of the second (high) range of the lever.

[0011] Although physical levers are described herein, computerized or automated controllers may also be implemented in accordance with embodiments herein. For example, an aircraft-wide system controller may request a particular thrust level, similar to how a physical thrust lever is moved to request a particular thrust level. The hybrid-electric powerplant described herein may have its own controller that may respond to such a request in the same manner that it responds to a physical lever. In other words, the aircraft-wide system controller may not request or provide information related to a particular power generation mode, but may instead request a desired level of thrust, and the hybrid-electric powerplant described herein may respond accordingly to provide different thrust levels across multiple operating modes without requesting a particular operating mode.

[0012] Advantageously, such an embodiment provides simplified operation of the hybrid-electric powerplant that can provide maximum flexibility and options for in-flight performance. Pilot training and the potential for pilot error are reduced. This one-lever system may also advantageously enable a simplified overall supervisory controller for the hybrid-electric powerplant, which uses deterministic coding techniques and provides a quicker, less burdensome path to Federal Aviation Administration (FAA) certification.

[0013] Hybrid generator set and its operating modes Aircraft typically include custom-designed propulsion mechanisms and methods for powering those propulsion mechanisms. In this way, the propulsion mechanisms and the power provided to those propulsion mechanisms can be optimized to provide the amount of propulsive force required for a particular type and size of aircraft, while minimizing the weight of components within the aircraft. In other words, the propulsion mechanisms and the power of those propulsion mechanisms are often optimized for a particular type and size of aircraft, so that components of one aircraft cannot be easily used with different types of aircraft drive architectures, such as direct drive aircraft, parallel drive aircraft, and serial drive aircraft.

[0014] Various embodiments of flexible architectures and optimized components thereof for aerospace hybrid systems are described herein. A hybrid system may be or include a system in which fuel is combusted in a piston, rotary, turbine, or other engine, and the output of the piston engine is operatively connected to a generator to output electrical power. The embodiments described herein include a flexible system capable of powering many different types of aircraft and propulsion mechanisms. Such a system may advantageously reduce the design complexity of different types of aircraft, reduce the manufacturing costs of such systems since less customization allows for economies of scale in mass production of the systems, and ultimately reduce the complexity of aircraft using the systems described herein.

[0015] The flexible architecture described herein may further be used to power propulsion mechanisms in different ways, either in the same aircraft or in different aircraft. For example, a flexible architecture for powering propulsion mechanisms may operate in a number of different modes to power different types of propulsion mechanisms. A first aircraft may utilize one, some, or all of the different modes in which the flexible architecture may operate. A second aircraft may utilize one, some, or all of the different modes, and the modes utilized by the second aircraft may be different than the modes utilized by the first aircraft.

[0016] Thus, different aircraft may utilize different modes of powering the propulsion mechanisms provided by the flexible architecture described herein. Although the use of the flexible architecture may be customized in this manner, the physical hardware of the flexible architecture may be adapted for use in different aircraft with minimal or no modification to the physical components of the flexible architecture described herein. Instead, the use of different modes in different aircraft may be achieved largely based on how the components of the flexible architecture are controlled with a processor or controller. Thus, computer readable instructions may also be stored in a memory operatively coupled to the processor or controller, such that when the instructions are executed by the processor or controller, the computing device including the processor or controller may control the various components of the flexible architecture described herein to utilize any mode of use desired for a particular implementation, aircraft, flight phase, etc.

[0017] Aircraft power generation and propulsion systems may utilize various cooling systems to ensure that various components of the aircraft are kept at safe temperatures for operation, and may also maintain the components within temperature ranges that they can operate more efficiently. Further, advantageous cooling systems are described herein that leverage various aspects of the hybrid architecture described herein to efficiently cool components of the flexible architecture that powers the propulsion mechanisms of the aircraft.

[0018] An aircraft that includes hardware that provides different power modes for the propulsion mechanisms may have various components for which it is desirable to provide cooling. Thus, a single cooling system that efficiently moves air to the various components that enable the different power modes can reduce the weight of the aircraft and the power consumption of the cooling system. Figures 1-8 and the accompanying discussion below relate specifically to an example flexible architecture for powering an aircraft's propulsion system, and Figures 9-21 and the accompanying discussion below relate to various embodiments of the cooling system of the example flexible architecture.

[0019] 1A illustrates an example of a flexible architecture 101 for an aerospace hybrid system in accordance with an illustrative embodiment. As described herein, the flexible architecture 101 can be efficiently used in a wide range of applications using a single hybrid generator system that can be adapted in multiple ways depending on the requirements of the aircraft and the flight phase (e.g., used in different modes).

[0020] The flexible architecture 101 of FIG. 1A is a hybrid generator including an engine 105, a clutch 115, a generator / motor 121, and a power shaft 111. As described further below, the flexible architecture 101 can be used to achieve a variety of different modes, as needed, depending on the requirements of a particular aircraft installation or a particular flight phase. The engine 105 can be a combustion engine, such as an internal combustion engine. The engine 105 can be more specifically either a piston internal combustion engine, a rotary engine, or a turbine engine. Such engines can use standard gasoline, jet fuel (e.g., Jet A, Jet A-1, Jet B fuel), diesel fuel, biofuel alternatives, and the like. In various embodiments, other types of engines can also be used, such as small engines (e.g., Rotax gasoline engines) in drone implementations.

[0021] As mentioned above, the engine 105 may be a piston combustion engine. A piston combustion engine may advantageously rotate an output rotor or shaft at revolutions per minute (RPM) that may be more desirable for direct output to power a generator and / or a propulsion mechanism (e.g., a propeller) than other engines. For example, a piston combustion engine may have an output on the order of thousands of RPM. For example, a piston combustion engine may have an output of 2200-2500 RPM, which may be a desirable RPM for a propeller. In particular, the propeller may be designed to have a size that produces a desired tip speed of the propeller based on the RPM output of the piston combustion engine (e.g., 2200-2500 RPM). Other types of engines, such as turbine engines, may output rotational power on the order of tens of thousands of RPM, much higher than a piston combustion engine. In another embodiment, a motor / generator may be driven at the higher RPM of the turbine engine to benefit efficiency, power output, or other important factors. In some embodiments, a gearbox is added between the output of the high rotation engine and the other components of FIG. 1A to reduce the output speed of the engine 105. However, adding a gearbox can also increase the weight of the system, which is undesirable in some embodiments. Piston combustion engines may also have an additional advantage in terms of noise compared to turbine engines. Turbine engines are typically noisier than piston combustion engines, and the noise experienced by humans from turbine engines is typically more unpleasant to the listener than the noise produced by piston combustion engines. Quieter engines may also be more valuable in urban or more densely populated environments where reduced noise is desired.

[0022] The engine 105 outputs rotational power to the clutch 115, which may be controlled to engage or disengage the power shaft 111. In other words, the power shaft 111 may be engaged with the rotational output of the engine 105 by the clutch 115, such that rotational power may be transferred between the output of the engine 105 and the power shaft 111. When the clutch 115 disengages the output of the engine 105 and the power shaft 111, the power shaft 111 may rotate independently of the output of the engine 105. The clutch 115 may be physically located between the engine 105 and the generator / motor 121, or may even contact opposite sides of the engine 105 and the generator / motor 121 to reduce the overall footprint of the flexible architecture.

[0023] The generator / motor 121 can also be engaged or disengaged with the power shaft 111. In other words, the generator / motor 121 can be controlled OFF such that the rotation of the power shaft 111 does not cause the generator / motor 121 to generate power. Similarly, the generator / motor 121 can also be controlled ON such that the rotation of the power shaft causes the generator / motor 121 to generate power. The generator / motor 121 is referred to as a generator / motor because it can function as either a generator or a motor. In various embodiments, the generator / motor 121 may be referred to as an electric machine, which can be a generator, an electric motor, or both.

[0024] The flexible architecture further includes power inputs and outputs (I / O) 125 connected to the generator / motor 121. As described further herein, the generator / motor 121 can generate power based on the rotation of the power shaft 111 output via the power I / O 125 or can receive power via the power I / O 125 and use it to drive the power shaft 111.

[0025] The generator / motor 121 may also function as a driver for the power shaft 111. Upon receiving power via the power I / O 125 from a battery or other form of electrical energy storage elsewhere in the system, the generator / motor 121 may provide a rotational force to drive the power shaft 111. This may occur as long as the generator / motor 121 is controlled to be switched on so that it is engaged with the power shaft 111. If the generator / motor 121 is controlled to be switched off so that it is not engaged with the power shaft 111, then the power shaft 111 may not be rotated by the generator / motor 121.

[0026] The power output from the power I / O 125 can be used to drive electric motors of electric propulsion mechanisms (such as propellers). The power output from the power I / O 125 can also be used to power and / or charge other devices of the aircraft or aerospace vehicle. For example, the power output from the power I / O 125 can be used to charge one or more batteries. The power output from the power I / O 125 can also be used to power other devices or accessories of the aircraft or aerospace vehicle. The power I / O 125 also has an input so that the power shaft 111 can be driven by any power received via the power I / O 125 (such as power from one or more batteries). The power generated by the generator / motor 121 can be alternating current (AC) power. The AC power can be converted to direct current (DC) power by power electronics (e.g., a rectifier or inverter) and output to a DC bus. The DC bus can be connected to a battery and / or an electric propulsion mechanism. In this manner, the electric propulsion mechanism may provide power via the DC bus. In various embodiments, the motors of the electric propulsion mechanisms use AC power, and therefore the DC power from the DC bus may be converted from DC power to AC power before use by the electric propulsion mechanisms (e.g., by an inverter).

[0027] Any rotation of the power shaft 111 itself, whether driven by the engine 105 or the generator / motor 121, can also be used to drive one or more propulsion mechanisms. For example, the rotation of the power shaft 111 may be used to directly drive a propeller or may be used to power an electric motor that drives the propulsion mechanism. The rotation of the power shaft 111 may also operably drive a gearbox that is operably connected to another component, such as one or more propellers, one or more rotors, or other rotating devices for various aircraft applications.

[0028] Accessory pad 131 may also be coupled to engine 105 and may include a low voltage direct current (DC) generator for power, separate from generator / motor 121 and power I / O 125, which may be configured for high voltage and high power I / O. In some embodiments, generator / motor 121 may also include two different windings and power I / O 125 includes two different outputs (e.g., high voltage and low voltage). An accessory power source may be associated with one of the outputs of power I / O 125 in addition to or instead of the output of accessory pad 131. Accessory pad 131 may be used to power devices or accessories of the aircraft or aerospace vehicle that do not require the high voltage or high current output that generator / motor 121 can output on power I / O 125. High voltage (HV) on an aircraft may be, for example, 400 volts (V) or 800V, but may be any value between 50V and 1200V. The low voltage (LV) on an aircraft can be 12V, 14V, 28V, or any voltage less than 50V.

[0029] FIG IB illustrates an additional example of a flexible architecture 150 for an aerospace hybrid system in accordance with an illustrative embodiment. In particular, the flexible architecture 150 of FIG IB includes several components that may be the same or similar to those described above with respect to FIG 1A, including an engine 155, a clutch 175, a power shaft 180, and / or a generator / motor 185. The flexible architecture 150 further illustrates an output of the engine 155 in the form of a crankshaft 160 that is rigidly connected to an output flange 165. The output flange 165 is rigidly connected to one side of the clutch 175 by bolts 170.

[0030] The clutch 175 may be configured to engage the power shaft 180 to convert rotational motion from the crankshaft 160 and the output flange 165 to the power shaft 180. The clutch 175 may be further configured to disengage the power shaft 180, allowing the power shaft 180 to rotate independently relative to the crankshaft 160 and the output flange 165. Additionally, FIG. 1B illustrates how the rotatable components of the flexible architecture 150 may all be aligned along a single axis 190. The rotatable components of FIG. 1A may similarly be aligned along a single axis, as shown in FIG. 1B. Additionally, the power shaft 180 may be a splined shaft that fits into an inner diameter opening of the clutch 175 and the generator / motor 185. Features other than splines, such as tapers, may also be used. In either case, the generator / motor 185 and / or the clutch 175 may be configured to match and connect to splines, tapers, or other features on the power shaft 180 so that the components can properly engage with one another.

[0031] Advantageously, the generator / motor 121 and / or the generator / motor 185 of FIG. 1B may be used as a starter for the engine 105 or the engine 155, respectively. In other words, the generator / motor 185 may be used to start the engine 155 by rotating the crankshaft 160 while engaging the clutch 175. Such a system may be advantageous, for example, when the generator / motor 185 is powered by a battery or other power source. Thus, the engine 155 may be a piston combustion engine as described herein. The engine 155 may not require a separate starter component, reducing the weight and complexity of the flexible architecture described herein.

[0032] FIG. 2A illustrates a block diagram depicting an aircraft control system 200 used with a flexible architecture 201 for an aerospace hybrid system in accordance with an exemplary embodiment. The aircraft control system 200 may be used to realize, for example, one or more of the various modes described below in which the flexible architecture described herein may be used. The flexible architecture 201 may be the same as, similar to, or have some or all of the components of the flexible architectures 101 and / or 150 of FIG. 1A and / or FIG. 1B. The aircraft control system 200 may include one or more processors or controllers 205 (hereinafter controller 205), memory 210, a main aircraft controller 220, an engine 230, a generator / motor 235, a clutch 240, a power I / O 245, an accessory pad 250, and one or more sensors 260. The connections in FIG. 2A illustrate control signal related connections between the components of the aircraft control system 200. Other connections not shown in FIG. 2A may exist between different aspects of the aircraft and / or aircraft control system 200 to provide power, such as high voltage (HV) or low voltage (LV) power for the aircraft.

[0033] The memory 210 may be a computer-readable medium configured to store instructions. Such instructions may be computer-executable code executed by the controller 205 to implement various methods and systems described herein, including various modes and combinations of those modes using the flexible architecture herein. The computer code may be written such that various ways of implementing different modes of the flexible architecture herein are implemented automatically, for example, based on various inputs indicative of a particular flight phase (e.g., landing, takeoff, cruise, etc.). In various embodiments, the computer code may be written to implement various modes herein based on inputs from a user or pilot of the aircraft or aerospace vehicle, or may be implemented based on a combination of user inputs and automatic implementations based on non-human inputs (e.g., inputs from sensors on or off the aircraft, based on a planned flight plan, etc.). The controller 205 may be powered by an aircraft or aerospace vehicle power source, such as the accessory pad 131, one or more batteries, an output of the power I / O 125, an aircraft power bus powered by any power source, and / or any other available power source.

[0034] The controller 205 may also communicate with each of the engine 230, the generator / motor 235, the clutch 240, the power I / O 245, the accessory pads 250, and / or the sensors 260. In this manner, the components of the flexible architecture may be controlled to achieve the various modes described herein. In various embodiments, the engine 230, the generator / motor 235, the clutch 240, the power I / O 245, and the accessory pads 250 may be similar or similarly named components as shown in and described above with respect to FIG. 1A. The power I / O 245 may also include pre-charge electronics to protect the electrical components of the flexible architecture, including, for example, a direct current (DC) bus, from excessive inrush current during startup, as described herein. For example, if a high voltage (HV) bus is 400V and a new component is connected to the HV bus at 0V, the instantaneous inrush current may be very high and may cause damage to the HV bus and / or the component. As a result, the precharge electronics allow the voltage of the component to slowly increase before being fully connected to the HV bus or other power source.

[0035] The sensors 260 may include various sensors for monitoring various components of the flexible architecture 201. Such sensors may include temperature sensors, tachometers, fluid pressure sensors, voltage sensors, current sensors, status sensors for determining, for example, the current state of the clutch 250, or other types of sensors. For example, voltage and / or current sensors may be used to inform the function and settings of the motor / generator, the selected state of the clutch, or adjustments of other components of the system. The status sensors may also indicate a particular mode in which the flexible architecture is being used, and the system may receive input (e.g., from the pilot, from an automatic flight controller) to change the system to a different state or mode for a particular phase of the upcoming flight. Other sensors may include a pitot tube to measure the airspeed of the aircraft, an altimeter to measure the altitude of the aircraft, and / or a global positioning system (GPS) or similar geographic location sensor to determine a position relative to the ground and / or known / mapped structures.

[0036] The components within the dashed lines of flexible architecture 201 in FIG. 2A may be associated with the flexible architecture described herein, while main aircraft controller 220 may be associated with a broader aircraft system. In other words, main aircraft controller 220 may control aspects of the aircraft other than flexible architecture 201, while controller 205 controls aspects of the aircraft related to flexible architecture 201. Main aircraft controller 220 and controller 205 communicate with each other to coordinate power supply to various propulsion mechanisms of the aircraft. For example, main aircraft controller 220 may send signals to controller 205 requesting a particular power output level of one or more particular propulsion mechanisms. Controller 205 can receive such control signals and, based on the control signals from main aircraft controller 220, determine how to adjust flexible architecture 201 to output the desired power level (e.g., which mode to enter, how to control elements of flexible architecture 201). In various embodiments, main aircraft controller 220 can send signals related to controlling particular aspects of flexible architecture 201. In other words, in addition to or instead of sending a desired power output signal to controller 205, controller 205 acts as a relay that retransmits control signals from main aircraft controller 220 to the components of flexible architecture 201, from which controller 205 determines how to control the individual components of flexible architecture 201.

[0037] In various embodiments, main aircraft controller 220 may also transmit control signals related to future desired power output, future flight phases, flight plan information, or the like. In this manner, controller 205 may receive and use information regarding the anticipated power demands of the aircraft to determine how to control aspects of flexible architecture 201, both present and future. For example, flight plan information may be used to determine when to use battery power, when to charge batteries, etc. In another example, if a large demand for power is anticipated, controller 205 may cause engines 230 to begin operating at a desired RPM to provide a desired level of power.

[0038] In various embodiments, controller 205 may also communicate with one or more batteries to monitor the charge levels of the batteries, control when the batteries are charged or discharged, control when the batteries are used to power generators / motors 235, and control when the batteries are used to directly power other aspects of the aircraft. However, in other embodiments, main aircraft controller 220 may communicate with the aircraft's batteries and / or relay information related to the batteries and their control to controller 205. Similarly, in cases where the aircraft's batteries are controlled by main aircraft controller 220 rather than controller 205, controller 205 may send control signals related to the batteries to the main aircraft controller, thereby controlling the batteries as needed or desired for the functioning of flexible architecture 201.

[0039] In various embodiments, the power I / O 245 can include two different outputs (e.g., a high voltage (HV) output and a low voltage (LV) output) associated with two different windings of the generator / motor 235. Thus, two different voltages (e.g., HV and LV) can be output and controlled by the controller 205 and / or the main aircraft controller 220. The power I / O 245 can additionally or alternatively include a voltage conversion component (e.g., a DC-DC converter) to output two or more different voltages. In such an embodiment, the two different outputs can be achieved without using two separate windings. The two different outputs can be output, for example, to different power buses on the aircraft (such as an HV bus and an LV bus). The two outputs of the power I / O 245 can also be controlled independently by the controller 205. Thus, the outputs can be turned off (e.g., by turning off the motor / generator field current to allow the generator power shaft and rotor to rotate or freewheel relative to the rest of the motor / generator).

[0040] In some embodiments, the accessory pads may not be controlled by controller 205 and / or main aircraft controller 220. The accessory pads may be on whenever engines 230 are running, or may be controlled separately (e.g., by a manual switch that a user toggles) to control when and how to power the aircraft accessories.

[0041] In some embodiments, controller 205 may be in communication with a wireless transceiver that may be onboard an aircraft or aerospace vehicle, thereby allowing controller 205 to communicate with other computing devices that are not hardwired to system 200. In this manner, instructions or inputs for implementing the various modes of the flexible architecture described herein may also be received wirelessly from a computing device of a remote device. In other embodiments, system 200 may only communicate with components onboard an aircraft.

[0042] Figure 2B illustrates a block diagram depicting a second aircraft control system 275 for use in the flexible architecture for an aerospace hybrid system in accordance with an illustrative embodiment. In the example of Figure 2B, system 275 does not have a separate main aircraft controller as in Figure 2A. Instead, the entire aircraft has a single main controller 280 that controls the flexible architecture and all aspects of the aircraft (e.g., including the aircraft's propulsion mechanism 255).

[0043] The controller 285 may communicate with one or more propulsion mechanisms 255 of the aircraft to control them. The controller 285 may also communicate with one or more sensors 270 of the aircraft or aerospace vehicle, which may be aircraft sensors and flexible architecture sensors. In particular, the sensors 260 may be embedded in any of the components of Figures 1A and / or 1B described above, and thus may be used to inform how to control the devices of Figures 1A and / or 1B and / or how the modes described herein are implemented as described herein.

[0044] 2A or 2B, controller 205, controller 285, and / or main aircraft controller 220 may also be in communication with a cooling system configured to cool and / or heat any component of the flexible architecture, one or more batteries, or other aspects of the aircraft, such that the cooling system may also be controlled in conjunction with other systems and methods described herein.

[0045] Described below are five particular modes that may be implemented using various embodiments of the flexible architecture described herein (e.g., including the flexible architectures shown in and described with respect to Figures 1A, 1B, 2A, and 2B).

[0046] In a first mode, which may be referred to herein as a hybrid generator mode, a clutch (e.g., clutch 115 of FIG. 1A and / or clutch 175 of FIG. 1B) may be controlled to engage an engine (e.g., engine 105 of FIG. 1A and / or engine 155 of FIG. 1B) with a power shaft (e.g., power shaft 111 of FIG. 1A and / or clutch output / power shaft 180) extending between the clutch and a generator / motor (e.g., generator / motor 121 of FIG. 1A and / or generator / motor 185 of FIG. 1B), such that the engine may rotate a power shaft within the generator / motor to generate electrical power that is provided to other systems (such as propulsion / systems) of the aircraft via a power I / O (e.g., power I / O 125 of FIG. 1A). For example, such propulsion / systems may be driven using an electric motor, and the electrical power output by the generator / motor in the first mode may be used to drive such propulsion / systems. In short, in the first mode, the engine can engage the power shaft using the clutch to drive the generator / motor and output electrical power from the generator / motor.

[0047] In a second mode, which may be referred to herein as a direct drive engine mode, a clutch (e.g., clutch 115 of FIG. 1 and / or clutch 175 of FIG. 1B) may engage the output of an engine (e.g., engine 105 of FIG. 1A and / or engine 155 of FIG. 1B) to a power shaft (e.g., power shaft 111 of FIG. 1A and / or clutch output / power shaft 180) that passes through a generator / motor (e.g., generator / motor 121 of FIG. 1A and / or generator / motor 185 of FIG. 1B) to provide mechanical power to a propulsion mechanism, such as a propeller of an aircraft. In such a mode, the magnetic field (electric field) is removed from the generator / motor (e.g., the generator / motor is controlled to be turned off or disengaged), causing the generator / motor power shaft and rotor to rotate or freewheel, and thus the generator / motor power I / O (e.g., power I / O 125 of FIG. 1A) is disengaged and no longer outputs power. That is, in the second mode, the engine drives the power shaft to mechanically or otherwise power the propulsion mechanism, while the power shaft rotates within the generator / motor but does not receive or output power at the power I / O.

[0048] In a third mode, which may be referred to herein as an augmented thrust mode, a clutch (e.g., clutch 115 of FIG. 1 and / or clutch 175 of FIG. 1B) may engage an engine (e.g., engine 105 of FIG. 1A and / or engine 155 of FIG. 1B) to a power shaft (e.g., power shaft 111 of FIG. 1A and / or clutch output / power shaft 180) that passes through a generator / motor (e.g., generator / motor 121 of FIG. 1A and / or generator / motor 185 of FIG. 1B), which is used as a motor to draw power from an external source, such as a battery pack, via power I / O (e.g., power I / O 125 of FIG. 1A). This provides a higher mechanical power output to the power shaft than the engine or generator / motor can provide. That is, in the third mode, both the engine and the generator / motor are used to simultaneously drive the power shaft to send power to the propulsion mechanism.

[0049] In a fourth mode, which may be referred to herein as a direct drive generator / motor mode, the clutch (e.g., clutch 115 of FIG. 1 and / or clutch 175 of FIG. 1B) may disengage the engine (e.g., engine 105 of FIG. 1A and / or engine 155 of FIG. 1B) from the generator / motor (e.g., generator / motor 121 of FIG. 1A and / or generator / motor 185 of FIG. 1B) and thereby provide power to the generator / motor via the power I / O (e.g., power I / O 125 of FIG. 1A) to drive the generator / motor as a motor and provide mechanical power to the power shaft (e.g., power shaft 111 of FIG. 1A and / or clutch output / power shaft 180). That is, in the fourth mode, only the generator / motor may provide power to the propulsion mechanism based on power received at the power I / O.

[0050] In a fifth mode, which may be referred to herein as a split engine power mode, a clutch (e.g., clutch 115 of FIG. 1 and / or clutch 175 of FIG. 1B) may engage an engine (e.g., engine 105 of FIG. 1A and / or engine 155 of FIG. 1B) to a generator / motor (e.g., generator / motor 121 of FIG. 1A and / or generator / motor 185 of FIG. 1B), which may cause the engine to rotate the generator / motor as a generator to provide electrical power to other systems of the aircraft via power I / O (e.g., power I / O 125 of FIG. 1A) and to provide mechanical power to a power shaft (e.g., power shaft 111 of FIG. 1A and / or clutch output / power shaft 180) to drive a system such as a propeller. That is, in the fifth mode, the engine may be used to drive a power shaft and generator / motor to output electrical power via the power I / O and power shaft.

[0051] As described herein, any of these five modes (or variations thereof) may be used in a single flexible architecture described herein. Furthermore, particular modes and / or combinations of modes may be beneficial for particular aircraft or aerospace vehicle types, particular propulsion types, particular flight phases of an aircraft or aerospace vehicle, etc.

[0052] For example, in a hybrid electric vertical take-off and landing (VTOL) aircraft that includes electric motor-driven propellers, the flexible architecture herein can be used only as a power source, such that the flexible architecture can power the aircraft in a first mode (e.g., hybrid generator mode) during any portion of the flight phase that requires powering the aircraft's power bus or one or more motors of the aircraft.

[0053] In another example, an aircraft including a single large main pusher propeller (e.g., at the rear of the aircraft fuselage) and an array of electric motors / propellers (e.g., on the wings of the aircraft) may use a flexible architecture in a fifth mode (e.g., split engine power mode) to mechanically power the main pusher propeller and electrically power the wing-mounted motors during takeoff. Figures 3 and 4 show two examples of such aircraft 300 and 400 that may use the flexible architecture for aerospace hybrid systems according to exemplary embodiments. For example, aircraft 300 has a main pusher propeller 305 and aircraft 400 has a main pusher propeller 405 in the form of a ducted pusher fan. In both examples, the fifth mode described herein may be used to provide mechanical power to the main pusher propellers 305 and 405 from a power shaft. Additionally, the wing-mounted electric motors / propellers 310 and 410 may be driven by electrical power from a motor / generator as described herein.

[0054] Alternatively, the flexible architecture described herein may be used to power the configuration as shown in Figures 3 and 4 in a third mode (e.g., augmented thrust mode) during takeoff, with the battery pack powering both the wing-mounted motors and augmenting engine power on the power shaft driving the main propulsion propeller. During cruise flight, the aircraft may use the second mode (e.g., direct drive engine mode) to drive only the main propulsion propeller. In another example, during cruise flight, the aircraft is equipped with a clutch between the power shaft and the propulsion propeller, and the controller can operate the aircraft in a first mode (e.g., hybrid generator mode) to drive the wing-mounted motor by disengaging the power shaft from the propulsion propeller and outputting power from the generator / motor to the wing-mounted motor. In another example (e.g., emergency situations such as engine failure), the propulsion propeller can be driven in a fourth mode (e.g., direct drive generator / motor mode) using power input to the power I / O, such as from one or more batteries.

[0055] In another example, the aircraft may be a VTOL aircraft including a gyrocopter-type main rotor that may operate with or without power, and may have forward thrust motors and propellers mounted on the wings. In one embodiment, the flexible architecture may be used entirely in a first mode (e.g., hybrid generator mode) where the power provided from the power input / output (and generator / motor) drives the motors coupled to the gyrocopter-type main rotor and the power is used to drive the motors mounted on the wings. In one embodiment, the aircraft is also configured to include a clutch between the power shaft and the gyrocopter-type main rotor, whereby the flexible architecture can rotate the gyrocopter-type main rotor (e.g., to get the gyrocopter-type rotor up to speed for takeoff) using a second mode (e.g., direct drive engine mode) or a third mode (e.g., augmented thrust mode). In such an example, the controller may then switch the flexible architecture to the first mode (e.g., hybrid generator mode) after the gyrocopter-type rotor has reached speed (e.g., switching to the first mode for cruise flight). The fourth mode (e.g., direct drive generator / motor mode) can be used again in the event of engine failure to drive the power shaft (and thus the gyrocopter-type rotor) using power from a power source such as one or more batteries.

[0056] 5 illustrates another example of an aircraft 500 in which a flexible architecture for an aerospace hybrid system according to an example embodiment may be used. For example, the aircraft 500 may include multiple (e.g., eight) electric motors / propellers 505 on tilt wings that may be powered using a first mode (e.g., hybrid generator mode) described herein, in which an engine may be engaged with a power shaft using a clutch to drive a generator / motor that may output power to the various electric motors / propellers 505 on the tilt wings.

[0057] Accordingly, an advantageous flexible architecture for an aircraft is described herein that can provide a variety of modes for powering the propulsion. While a particular aircraft and propulsion configuration may not utilize each of the modes described herein that the flexible architecture can perform, the flexible architecture can be implemented in a variety of aircraft to provide a variety of modes. Similarly, while an example flexible architecture that includes five different modes for powering the propulsion is described herein in detail, other flexible architectures that have fewer, more, or different modes for powering the propulsion are contemplated herein.

[0058] For example, the flexible architecture may not have a clutch as described herein and still implement the various modes described herein where it is desirable to couple the engine output to the motor / generator and / or output power shaft of the system. For example, in a first mode, the engine may rotate a power shaft to cause a generator to generate electricity. In a second mode, the engine may directly drive, for example, mechanical propulsion components, but it is not necessary to disengage the engine from the motor / generator or power shaft because the motor / generator may be turned off or the motor / generator power shaft and rotor may be allowed to freewheel within the motor / generator. In a third mode, the engine and motor / generator are used to drive the power shaft, so it is not desirable to use a clutch to disengage the engine and motor / generator. In a fifth mode, the engine may rotate a power shaft to cause a generator to generate electricity, such that the power shaft mechanically powers the propulsion mechanism. As such, in an aircraft using any of the first, second, third, and fifth modes described above, it is not necessary to disengage the power shaft from the engine output. Thus, in embodiments using any combination of the first, second, third, and / or fifth modes (but not the fourth mode), the system may not use a clutch because the engine output is constantly connected to the motor / generator power shaft, which may be beneficial since clutches can be heavy and unreliable.

[0059] 6 is a flow chart illustrating an example of a first method 300 for using a flexible architecture for an aerospace hybrid system during different flight phases of an aircraft including a main propulsion propeller, according to an exemplary embodiment. In particular, the aircraft may be an aircraft including a single larger propulsion propeller and an array of electric motors on the wings and corresponding smaller propellers. During the takeoff flight phase at 602, the fifth mode described herein may be used to mechanically power the main propulsion propeller and to electrically power the wing mounted motors. During the cruise flight phase at 604, the second mode described herein may be used to mechanically power only the main propulsion propeller and not to electrically power the small electric motor / propeller.

[0060] FIG. 7 is a flow chart illustrating an example of a second method 400 for using a flexible architecture for an aerospace hybrid system during different flight phases of an aircraft including a main propulsion propeller, according to an exemplary embodiment. In particular, the aircraft may be an aircraft including a single large propulsion propeller and an array of electric motors and corresponding smaller propellers on the wings. During the takeoff flight phase at 702, a third mode described herein, called augmented thrust, may be used to power the main propulsion propeller via a generator / motor (drawing power from the battery) and to mechanically power the main propulsion propeller directly from the engine. Additionally, during takeoff, power (generated by the generator / motor and / or directly from the battery) may also be provided to the electric motors on the wings. During the cruise flight phase at 704, a second mode described herein may be used to mechanically power only the main propulsion propeller and not the smaller electric motors / propellers.

[0061] Returning to FIG. 1A, when clutch 115 is engaged and engine 105 is powering power shaft 111 and generator / motor 121 is not running or turned on, power shaft 111 may freewheel within generator / motor 121 (e.g., second mode described above). Similarly, power shaft 180 of FIG. 1B may freewheel within generator / motor 185 in various embodiments. However, engine 105 and / or engine 155 may generate torque pulses in power shaft 111 and / or power shaft 180 when clutch 115 and / or clutch 175 are engaged with their respective power shafts 111 and / or 180, which torque pulses may be dangerous to generators such as generator / motor 121 and / or generator / motor 185. In other words, large torque pulses on the shafts, similar to those that may occur when certain types of engines (e.g., diesel piston combustion engines) fire, can cause high angular accelerations that can cause fatigue or damage to components of generator / motor 121 and / or generator / motor 185 that are coupled to power shafts 111 and / or 180. Therefore, components such as flywheels or other heavy damping or spring coupling systems can be used to smooth out the torque on power shafts 111 and / or 180.

[0062] FIG. 8 illustrates an example of a flexible architecture 800 for an aerospace hybrid system having a flywheel for absorbing vibration torque, according to an exemplary embodiment. In particular, the flexible architecture 800 includes similar or identical components as shown in and described with respect to FIG. 1B, but includes a flywheel 195 rigidly connected to an output flange 165 by bolts 170. The flywheel 195 is further rigidly connected to one side of a clutch 175 by bolts 198. Thus, rotational motion can be transferred from the engine 155 through the crankshaft 160, the output flange 165, and the flywheel 195 to the clutch 175. The clutch 175 can then engage or disengage the power shaft 180 to selectively transfer the rotational motion received from the flywheel 195 to the power shaft 180. The flywheel 195 may further be, for example, a dual mass flywheel or a spring coupling.

[0063] In various other embodiments, a flywheel may not be used. For example, further embodiments of damping systems and devices are described herein that can dampen torque on a power shaft (e.g., power shaft 111) but do not include a flywheel. Additionally, in various embodiments, a flywheel may be used in combination with other damping systems or components to dampen or smooth the torque applied to the power shaft.

[0064] For example, a power shaft or rotor within the generator / motor itself may be rigidly coupled to the crankshaft of the generator / motor. In this way, the crankshaft and rotor together may dampen torque pulses on the power shaft or rotor, reducing tangential acceleration due to torque pulses from the engine. In such an embodiment, a clutch may be omitted. As such, the damping system is internal to the generator / motor, and the footprint and weight of the damping system may be smaller than a flywheel or other damping system that may be external to the generator / motor. In particular, the rigid coupling of the power shaft or rotor to the crankshaft increases the inertia of the power shaft or rotor, whereby the additional inertia helps to prevent the power shaft from rotating in a manner that slows it down or makes it more susceptible to acceleration due to torque pulses from the engine. In such an embodiment, the power shaft or rotor and crankshaft may function similarly to a flywheel.

[0065] In various embodiments, a generator / motor may be used that has a static inner portion and a rotating outer portion, which increases the inertia of the rotating portion and prevents the magnets in the generator / motor from rotating and becoming dislodged by the torque spike. In other words, the magnets are already rotating in the outer portion and therefore a constant steady radial force may be applied in addition to the tangential inertial force due to the acceleration of the torque spike.

[0066] The torque damping system may also be configured as part of the power shaft or rotor that connects the engine output to the generator / motor. For example, the hub between the generator / motor power shaft or rotor may include a coupling that includes a torsional spring and / or damping properties. A torsional damping coupling may include an elastomeric part or spring (e.g., made of steel or other metal) that reduces potentially harmful torque impulses from being passed from the engine output to the generator power shaft or rotor. A torsional damping coupling may be similar to or referred to as a resonant damping coupling. For example, such a torsional damping coupling may reduce the weight and size of the overall system as opposed to a system that uses a flywheel or other bulky damping system. One or more torsional damping couplings may be installed in the engine, between the engine and the clutch, in the clutch, between the clutch and the generator, and / or in the generator to achieve damping before the power shaft or rotor damages components of the generator itself.

[0067] Other methods of damping the torque of the generator power shaft or rotor may also be used. For example, the generator magnetic field may be controlled to generate pulses that act on the generator power shaft or rotor to partially or completely counteract the torque pulses imparted to the power shaft or rotor by the engine. Such pulses of the generator magnetic field may be controlled based on measurements of torque pulses applied by the engine, thereby preventing damage to generator components by the diesel engine. For example, in the third mode described above, where both the engine and the generator / motor provide power to the power shaft, pulses from the generator may be applied to the power shaft to simultaneously provide power to the power shaft and prevent damage to the generator components. In other modes described herein, the generator may be used to apply pulses to the power shaft whenever the power shaft is driven in part or in whole by the engine. Thus, to adequately protect the generator components in this manner, the pulses applied to the power shaft or rotor by the generator magnetic field may be configured to correlate with the torque pulses of the engine so that they can be appropriately counteracted.

[0068] Examples of how the flexible architectures described herein can be packaged and / or used in actual aircraft are further described below. For example, certain aircraft may use electric motors to drive their propulsion systems and therefore must include sufficient on-board electrical energy to drive those propulsion systems or a method for generating such on-board electrical energy. Additionally, regulations in a given jurisdiction may also require sufficient reserve energy to comply with aircraft operating regulations. The flexible architectures described herein can provide electrical energy and / or reserve energy for such propulsion systems, thereby enabling the systems described herein to operate with a variety of electric aircraft. For example, embodiments herein can efficiently convert jet fuel (or other liquid or gaseous fuels) to electricity, thereby enabling electric aircraft to be powered using widely available fuel sources.

[0069] FIG. 9 illustrates a perspective view of an example of a flexible architecture 900 of an aerospace hybrid system in accordance with an exemplary embodiment. This hybrid unit can be used as a core power plant for a variety of aircraft types and implementations. The hybrid unit of FIG. 9 is a tightly integrated power plant that can include some, all, and / or additional elements shown and described with respect to FIG. 1A, FIG. 1B, FIG. 2A, FIG. 2B, and / or FIG. 8.

[0070] Additionally, the hybrid unit may include an integrated cooling system 905 for cooling various aspects of the hybrid unit, heat exchangers associated with the hybrid unit, or heat sinks such as finned attachments for any aspect of the hybrid unit. The power output 910 may be or be connected to a power shaft (e.g., power shaft 110 of FIG. 1A, power shaft 180 of FIG. 1B or FIG. 8), thereby outputting rotational power from the hybrid unit to a propulsion system or other aspects of the aircraft. The electrical connector 915 may also be used to output electrical power (or input electrical power) as described herein. The electrical connector 915 may be, for example, an Amphenol Surlok Plus® connector or equivalent, or any other type of suitable connector. In this manner, a main bus, such as a direct current (DC) bus of the hybrid unit, may be connected via the electrical connector 915 (e.g., power input / output 125 of FIG. 1, power I / O 245 of FIG. 2A or FIG. 2B). These or other connectors may also facilitate connection and control of the components of the hybrid unit using a Controller Area Network (CAN) bus, a CAN 2.0 bus, and / or an SAE J1939 bus, etc. Such communication buses may operate at various speeds, such as 250 kilobytes per second (kbps), 500 kbps, 1000 kbps, etc. In various embodiments, the electrical connector 915 and / or other connectors may be customized for specific applications, such as various types of aircraft and the communications and power systems used by those aircraft.

[0071] By virtue of the power output 910 and the electrical connector 915, the hybrid unit of FIG. 9 may output mechanical power via the power output 910 and / or output electrical power via the electrical connector 915 and a DC bus within the hybrid unit (e.g., the power input / output 125 of FIG. 1 , the electrical I / O power 245 of FIG. 2A or FIG. 2B ). Similarly, just as mechanical power is received via the power output 910 to generate electricity for output via the electrical connector 915, electrical power may be received via the electrical connector 915 to drive the power output 910. For example, if the aircraft includes one or more batteries, additional power from the batteries may be received via the electrical connector 915 to augment the power applied to the power output 910, such that the power output 910 is driven by both the engines and electrical power from the aircraft's batteries, as described herein.

[0072] The hybrid unit of FIG. 9 may further include a connector 925 for connecting the engine to a fuel source. The connector 925 may be a quick fuel connector, such as an AN6 quick fuel connector. In this manner, the engine may be fueled to power the power output 910 and / or generate electricity that is output via the electrical connector 915. The hybrid unit of FIG. 9 may further include mounting hardware 920 for mounting the hybrid unit to an aircraft. In FIG. 9, the mounting hardware 920 is shown on the top of the hybrid unit, but in other embodiments, the mounting hardware may additionally or alternatively be located on either the top, bottom, side, etc. of the hybrid unit such that the hybrid unit may be mounted to the aircraft as desired.

[0073] Figure 10 illustrates a top view 1000 of the example flexible architecture of Figure 9 in accordance with an exemplary embodiment. Figure 11 illustrates a side view 1100 of the example flexible architecture of Figure 9 in accordance with an exemplary embodiment.

[0074] Thus, the hybrid units described herein can be used to power electric or hybrid electric aircraft and can provide better power than battery packs alone. For example, the hybrid units shown in Figures 9-11 can provide better energy density than batteries (e.g., 5-7 times better energy density). For example, the hybrid units described herein can have an energy density equivalent to 600-1200 watt-hours per kilogram (Wh / kg). The hybrid units described herein can also have the advantage of being more fuel efficient than other systems (e.g., 40% more fuel efficient than turbine engines) and can use readily available fuels such as Jet A, diesel, kerosene, biofuel alternatives, or other suitable or desired fuels. In other words, the hybrid units described herein can include engines, generators, inverters, and thermal management using air cooling in a compact package, so that aircraft with flexible architectures can advantageously utilize these components as a power plant. Various voltage outputs (e.g., 400 volts (V), 800V, 1000V, 1200V, etc.) are provided from the hybrid architecture, with connections for other accessories or system power (e.g., 28V). The flexible architecture described herein can be quieter than other systems (e.g., quieter than turbine engine systems). For example, within 100 feet of current systems, the noise level can be less than 70 decibels (dB).

[0075] The flexible architecture described herein is also scalable. For example, a large aircraft may use two or more of the flexible architectures described herein. The flexible architecture may also be used in different aircraft designed for different functions and purposes. For example, the flexible architecture described herein may be useful in urban air mobility (UAM) systems such as electric vertical take-off and landing (eVTOL) aircraft, electric short take-off and landing (eSTOL) aircraft, and electric conventional take-off and landing (eCTOL) aircraft. An example of a flexible architecture as shown in Figures 9-11 may have the specifications shown in Table 1 below. [Table 1] *Maximum burst shaft power depends on battery configuration. **Dry mass includes engine, generator, inverter, and thermal system.

[0076] As shown above, a 185 kW hybrid unit may be provided. Thus, two hybrid units may be provided on a given aircraft to provide 370 kW of power.

[0077] Figure 12 illustrates a perspective view 1200 of another example of a flexible architecture for an aerospace hybrid system in accordance with an illustrative embodiment. The flexible architecture of Figure 12 includes an engine 1205 and a generator, although the generator is hidden or not visible by other components of the system, such as cooling ducts. However, similar to the hybrid units of Figures 9-11, mechanical output power 1210 and electrical output power 1220 (both of which can also optionally receive electrical power) are provided.

[0078] In this manner, various embodiments herein provide a hybrid electric powerplant that can be incorporated into a variety of different types of aircraft in the aerospace market, thereby eliminating the need for aircraft manufacturers to build their own systems consisting of engines, generators, power electronics, cooling systems, and / or control systems to power their aircraft, which can be advantageous because the development process to create a powerplant system and certify it to aerospace standards can take four or more years and cost in excess of $10 million.

[0079] In this manner, the hybrid powerplant or flexible architecture described herein may be designed, manufactured, etc., separately from the design of the aircraft. Some aspects of the flexible architecture may be customized according to the desires of the aircraft manufacturer, but in a manner that does not result in a redesign or reconfiguration of the entire system. Thus, embodiments herein provide an integrated unit that includes an engine, generator, power electronics, cooling system, and / or control system in one package that is installed on the aircraft. Combining these elements into a single standalone unit may further advantageously allow the unit to go through the Federal Aviation Administration (FAA) certification process as a system. The certification system may then be used by multiple aircraft manufacturers, thus reducing the certification burden and development burden on aircraft developers, as well as improving efficiency by eliminating the need for multiple aircraft manufacturers to certify many different powerplant systems that are specifically designed for their aircraft.

[0080] By providing a combined unit that includes the engine, generator, power electronics, cooling system, and / or control system, the hybrid flexible architecture described herein can be optimized as a whole system rather than as individual components, and not as part optimized. Furthermore, such hybrid units can be used in multiple aircraft designs, but the systems designed as part of the aircraft design process are configured in a way that makes them difficult to re-apply elsewhere. Having a hybrid unit that can be applied to multiple market segments and aircraft designs that have common power requirements expedites aircraft development when the aircraft's main components (such as the hybrid unit or flexible architecture) are already certified and produced.

[0081] Hybrid electric systems for aviation have historically been designed from the ground up for each application / aircraft. Such a process is inefficient and is addressed by embodiments herein. For example, some aircraft have unique power plants designed specifically for the aircraft. Such solutions may include custom engines, generators, power electronics, control systems, cooling systems, battery packs, propulsion motors, and / or propellers. Embodiments herein provide a compact hybrid system for aircraft that constitutes two distinct halves within the aircraft's power and propulsion system, i.e., the upstream and downstream halves of a powertrain (such as the hybrid powertrain described herein).

[0082] 13 illustrates example downstream and upstream components for propelling an aircraft 1300 in accordance with an example embodiment. For example, downstream components 1310 of the aircraft system may include motors, rotors / propellers, attitude control components, etc. that are further related to the particular design of the aircraft. Upstream components 1305 of the aircraft that may be reused in different aircraft may include engines, generators, batteries, power distribution, fuel, generator noise reduction, etc.

[0083] Specifically, the upstream end of the powertrain may include hybrid powertrain elements responsible for generating electrical power. Such upstream components 1305 may include engines, generators, power electronics, control systems (for the upstream power generating components), cooling systems (for the upstream components), battery packs, and / or fuel. The downstream end of the powertrain may include hybrid powertrain elements that convert electrical power into thrust, attitude control, and / or active control of aerodynamics. These downstream components 1310 may further include electric motors, propellers, motor controllers, and / or control systems for the propulsion system.

[0084] Thus, there may be a need for a common upstream powertrain between very different electric aircraft designs with similar size and total power requirements. However, because downstream powertrains have little consistency between aircraft, these components may not be standardized to work across as many aircraft designs as the upstream components. Furthermore, upstream elements suitable for standardization may include components that are associated with power requirements but not with total energy requirements. In the case of engines, generators, power electronics, cooling systems, and / or control systems, these elements of the upstream powertrain may be sized to the specific power requirements (kW or hp) of the aircraft. However, the amount of fuel and the size of the battery packs are determined by the total energy requirements (kWh or hp hr), which may vary from aircraft to aircraft. In such an embodiment, the amount of fuel may be increased or decreased by changing the size of the fuel tanks to meet the design requirements of the aircraft, and the capacity of the battery packs (kWh) may be increased or decreased by adjusting the number of parallel stacks of cells in the battery packs or by adding additional battery packs.

[0085] Accordingly, embodiments are provided herein that provide a hybrid powerplant that tightly integrates an engine, generator, power electronics, control system (for the power generation system), and / or cooling system in a weight- and space-efficient manner, and that can be certified as a stand-alone unit that is separable from the aircraft and designed to provide propulsion.

[0086] Additionally, as described herein, the rotor in the generator can be optimized to serve multiple purposes in the context of a hybrid power plant. A conventional combustion engine may have a flywheel mass attached to the rotating shaft to improve smoothness of operation. However, in the context of an aerospace system, adding extra mass may not be attractive. As described herein, when coupling an engine to a generator in a hybrid power plant, the rotor of the generator can be designed to withstand torque impulses from the engine and be the rotating mass that the engine utilizes to smooth operation.

[0087] Additionally, while auxiliary power units are known in the art, these systems may be designed for a different purpose than as the primary source of propulsion for an aircraft, and therefore may not have a control system that is certifiable to the standards required for use in propulsion. Additionally, such systems may be designed without a cooling system, the aspect of which is left to the airframe designer. As such, these systems are not certified under Part 33 (FAA regulations for aircraft power plants). Additionally, these auxiliary power unit systems are designed as light-weight auxiliary systems that are used intermittently, rather than for high-efficiency propulsion systems used during all phases of flight. Additionally, while auxiliary power units may be designed to generate alternating current (AC) power, the hybrid-electric power plants described herein can generate direct current (DC) power, and the battery packs provide DC power and are charged using DC power, so that the hybrid-electric power plants can be coupled with large propulsion battery packs.

[0088] Turbogenerators are one type of adaptive auxiliary power unit proposed for hybrid power. Such systems lack cooling system integration that provides airframe developers with a cooling system that is part of the hybrid power plant. As such, airframe developers may have to design their own cooling system associated with the use of turbogenerators. Using the embodiments herein, an advantage is that a separate cooling system for cooling the hybrid power plant described herein does not have to be designed or developed for a particular airframe, because such a cooling system is already included in the flexible architecture described herein.

[0089] As such, the flexible architectures and hybrid electric powerplants described herein advantageously provide an engine for converting liquid fuel (or gaseous fuel) into rotary mechanical power, a generator coupled to the engine configured to convert the rotary mechanical power into electricity, and / or power electronics coupled to the generator and configured to convert a direct AC output of the generator into high voltage DC power. The flexible architectures and hybrid electric powerplants described herein further advantageously provide a control system configured to vary the power output of the engine to match the power demands of the aircraft's main propulsion electric buses to meet the power demands of the aircraft.

[0090] The hybrid power plant control system, power electronics, generator, and / or engine designs described herein may be adapted to meet regulatory requirements for reliability of propulsion aerospace systems (e.g., probability of failure is 10 -6 or less than 10 to the power of minus six). The flexible architectures and hybrid-electric powerplants may further include a control interface that enables the flexible architectures or hybrid powerplants to communicate with a vehicle-level flight control system to provide propulsion commands from the vehicle-level flight control system to the hybrid powerplant control system, and the control interface advantageously also enables the hybrid powerplant control system to send status messages (e.g., feedback for use in controlling the flexible architecture or hybrid powerplant) back to the vehicle-level flight control system. The flexible architectures and hybrid-electric powerplants may further include a cooling system that maintains a temperature range of the generator, power electronics, and / or engine over the full range of operational power output of the flexible architectures and hybrid-electric powerplants described herein.

[0091] Various embodiments of the flexible architecture or hybrid-electric powerplants described herein may further include a control system that varies the power output by varying engine torque and / or maintains a substantially constant revolutions per minute (RPM) over a significant range of power output. Such embodiments may speed up the response of the flexible architecture or hybrid-electric powerplant by eliminating throttle lag and increasing response times associated with the rotational inertia of the system.

[0092] Various embodiments of the flexible architecture or hybrid electric powerplant described herein may further include an option to provide a portion of the engine's power output as mechanical shaft power and a portion as DC power. Various embodiments of the flexible architecture or hybrid electric powerplant described herein may further include that the engine may be a piston engine, a diesel piston engine, a turbine engine, a rotary engine, or other type of combustion engine. Various embodiments of the flexible architecture or hybrid electric powerplant described herein may further include an example where the rotor of the generator is designed to be the flywheel of the engine. Various embodiments of the flexible architecture or hybrid electric powerplant described herein may further include a clutch between the engine and the generator to allow the generator to operate as a motor that can operate while the engine is shut down in some types of parallel hybrid installations described herein.

[0093] DC Bus Components Various embodiments are described herein for implementing a hybrid electric aircraft. Such aircraft may utilize a high voltage electric bus to distribute power to various components of the aircraft, such as motors for the propulsion of the aircraft. In such hybrid electric aircraft, it may be desirable to stabilize the high voltage electric bus within a certain predetermined voltage range (e.g., near a nominal voltage level) so that the propulsion motors can operate properly. Since various embodiments described herein may utilize a direct current (DC) bus, in particular, it may be desirable to maintain a desired DC voltage range. Advantageously, various embodiments herein efficiently maintain a desired DC voltage range on the DC bus by directly connecting at least one battery or supercapacitor to the DC bus, and further enable maintaining sufficient charge on the at least one battery or supercapacitor to maintain the desired DC voltage range on the DC bus. Such embodiments may prevent voltage spikes that may damage components of the hybrid electric or electric aircraft (e.g., electric motors and inverters for propulsion) and avoid voltage spikes or voltage sags that may adversely affect the reliability and / or performance and safety of the aircraft or systems of the aircraft.

[0094] In an electric aircraft, various embodiments of the overall architecture may include one or more power generating devices (e.g., generators) connected to a high voltage DC bus via a low impedance connection and providing power and energy to the bus. Within the same vehicle, one or more power consuming devices (e.g., electric motors) may be connected to the same DC bus that receive power and energy from the DC bus. Various embodiments of the electric aircraft may also include energy storage devices, such as battery packs or capacitors (e.g., supercapacitors), that can receive or provide power as needed depending on the bus voltage and the battery pack voltage.

[0095] For example, if a high voltage generator is generating DC power directly or is operating through a passive rectifier, the DC voltage generated by the motor may be a linear function of the motor revolutions per minute (RPM) of the shaft that rotates the generator. For example, a permanent magnet electric motor may generate a voltage based on the rotational speed (RPM). In many applications, the coupling of voltage and RPM may create motor control issues that limit the value of that electric motor in the system. To get more utility from a brushless motor without permanent magnets, an external voltage reference may be used to maintain a desired voltage level. A particular problem in aviation is the need for flight safety to precisely control power consumers (electric motors driving fans, propellers, or other devices) over a wide range of flight conditions that may not match the characteristics of the contributors (such as electric brushless generators). If the high voltage generator used is rotating slower than expected for any reason, the bus voltage may be lower than desired and motors on that bus may perform less than expected, leading to unsafe or undesirable conditions. If such a high voltage generator is spinning faster than expected, the bus voltage will be high and the motor performance may again deviate from expected or target values. Therefore, in applications of generators and motors sharing a common bus, it may be desirable to design the generators and motors used accordingly. In the case of electric aircraft, it is desirable to precisely control the motors to provide aircraft lift, thrust, aircraft attitude, etc. Therefore, compared to other non-aviation related implementations, it is desirable to better control the power provided to the motors (e.g., via the DC bus) by maintaining the power provided to the motors at a voltage that keeps the motors operating at a desired performance level. Furthermore, the power provided to the motors may be rapidly adjustable (e.g., providing the pilot or control system with a flexible and wide range over which the motor can be controlled) so that the pilot or control system of the aircraft can control the motor over a wide range as needed. In various embodiments, an inverter may be used to adjust the output voltage of an upstream generator, which is then used to provide power to the high voltage bus.The inverter can also be used to precisely control a downstream motor under varying load conditions.

[0096] Inverters allow system designers to extend the operating range of any motor and / or generator by controlling the current. For these inverters to function properly, it is advantageous to set and maintain the bus voltage that powers the inverters by some means other than the motor RPM (as it can be difficult to precisely control the voltage on the bus if only the motor RPM is used). Maintaining bus voltage is related to the expected variations in capacitance and load that are present under all system operating conditions. For example, a high voltage bus and power electronic system can become unstable if the load on that bus changes suddenly or if the capacitance (which acts like inertia in similar mechanical systems) is too low.

[0097] In various embodiments, the bus voltage can be established and maintained using a battery pack, a capacitor, or any combination thereof. Such devices can add capacitance and / or electrical inertia to the bus and may be passive, meaning that their intended function is governed entirely by the laws of physics, and may not require control or intervention (e.g., by a controller or control system). Supercapacitors (or ultracapacitors) have the additional desirable characteristic of high capacitance, but typically do not have a large energy storage capability. Supercapacitors can respond to very rapid fluctuations with very large power (e.g., energy over time). That is, they can provide stability to the bus for fluctuations that are relatively short in duration, low in amplitude, or the product of these two values ​​is relatively low. Batteries may be desirable because they have large capacitance for bus stability and can also store high energy. Batteries may not be able to respond to voltage changes as quickly as supercapacitors. This is because batteries are often rate limited in power application, especially when charging (discharge power capacity is often 10 times or more than charge capacity). For example, when current needs to be drawn from the bus to maintain a desired voltage level (e.g., when charging a battery), the battery may not be able to absorb that current as quickly as desired in certain embodiments (depending on the particular characteristics of the battery selected), although in some embodiments one or more battery packs may be sufficient to maintain the desired voltage level on the bus.

[0098] Accordingly, various embodiments are described herein that allow for independent control of one or more upstream generators and downstream motors with the addition of appropriately designed battery packs and / or supercapacitor banks to maintain a desired voltage on the DC bus. In an architecture where the voltage and capacitance of these storage elements are directly electrically connected to the main motor control elements on the bus (and not shielded by other switches, chargers, or similar devices), the battery packs and / or supercapacitor banks provide a lightweight and effective anchor or set point for the high voltage DC bus.

[0099] Aircraft battery packs may be deployed with hybrid electric power generation systems to support system safety standards applicable to flight articles. These battery packs and / or supercapacitors not only provide the required power or energy, but if selected to be set to the correct or desired voltage and connected to a high voltage motor controller, the battery packs and / or supercapacitor banks may provide a second valuable benefit of bus stabilization by connecting the battery packs and / or supercapacitor banks directly to the DC bus. The battery packs and / or supercapacitor banks may also be advantageously selected to have a target voltage for a given aircraft, although the actual voltage of the bus may of course vary somewhat depending on the state of charge (SOC) and changing electrical loads. The battery packs and / or supercapacitor banks may be advantageously selected to ensure that the actual voltage does not fall outside of a desired range. If the actual voltage falls outside of the desired range, or is expected to fall outside of the desired range, the aircraft controller or a hybrid generator set on the aircraft may adjust the power (e.g., torque) provided to the generator to increase or decrease the power provided to the DC bus to maintain the voltage within the appropriate desired range. Additionally, the RPM may be maintained at a constant or relatively constant level or within a predetermined range. Thus, the power provided to the generator or output to the power shaft may be adjusted by adjusting the torque output by the engine rather than adjusting the RPM of the engine's output. Additionally, it may be desirable to maintain the actual voltage set point to a range that may vary while remaining within a desired tolerance range for operating the electric motor or other aircraft components. Additionally, the battery pack may advantageously function as an auxiliary power source to drive the aircraft motor or other aircraft components in the event of a failure of the hybrid generator set generator or other components. This provides an increased level of safety and fault tolerance for the system.

[0100] 14 is a schematic diagram of an example system 1460 for providing a stable voltage to a direct current (DC) bus, according to an example embodiment. The system 1460 includes a hybrid generator set 1461 including a controller 1462, an engine 1463 connected by a shaft 1464 to a generator 1465, an inverter 1466, and a direct current (DC) bus 1467. The engine 1463 provides mechanical (e.g., rotational) power to the generator 1465 via the shaft 1464, which allows the generator 1465 to generate electrical power (e.g., alternating current (AC) power). The AC power from the generator 1465 is converted to DC power by the inverter 1466 and provided to the DC bus 1467. The inverter 1466 can also convert the AC power from the DC bus 1467 to AC power that the generator 1465 can use to provide a power output to the shaft (e.g., when the generator 1465 functions as a motor to provide power to aircraft components such as propulsion). The controller 1462 may control any component of the hybrid gen set 1461 (e.g., control the RPM output to the generator 1465). The controller 1462 may also measure characteristics of the DC bus 1467, such as the voltage of the DC bus and / or the current flowing through the DC bus 1467.

[0101] System 1460 also includes aircraft components such as inverters 1472 and 1476 connected to DC bus 1467, electric motors 1474 and 1478 connected to inverters 1472 and 1476, controller 1480, and battery packs 1482 and 1484. In various embodiments, aircraft components may include supercapacitors instead of or in addition to battery packs 1482 and 1484. In various embodiments, regardless of whether the aircraft components include separate batteries and / or supercapacitors, one or more battery packs and / or supercapacitors may be included as part of hybrid generator set 1461 and connected directly to a DC bus within hybrid generator set 1461. Although FIG. 14 shows multiple connections from DC bus 1467 of hybrid generator set 1461 to aircraft component 1470, other configurations are contemplated herein, such as a single connection to another bus of aircraft component 1470, or where DC bus 1467 is itself part of aircraft component 1470. The controller 1480 may communicate with the control 1462. In this manner, the controller 1480 may send information to the controller 1462 regarding how the inverters 1472 and 1476, the electric motors 1474 and 1478 are currently being controlled / used, or how the controller plans to use these components in the future. The controller 1480 may also monitor and measure the status of the battery packs 1482 and 1484 and send information related to their status (e.g., measurements related to state of charge, voltage, current flowing in or out of the batteries, etc.) to the controller 1462. In embodiments in which the hybrid generator set 1461 includes a battery or supercapacitor, the controller 1462 may monitor such components for similar information.

[0102] In various embodiments, an aircraft may include fewer, additional, or different elements than those shown in FIG.

[0103] FIG. 15 is a flow chart illustrating an example of a method 1500 for maintaining a stable DC bus voltage based on a communication from an aircraft-level controller, according to an exemplary embodiment. In operation 1502, a controller (e.g., controller 1462 of FIG. 14 ) may receive a communication from an aircraft controller (e.g., controller 1480 of FIG. 14 ) including power consumption or battery status information. The power consumption information may relate to how power is currently used, for example, by an aircraft inverter or electric motor. The power consumption information may relate to how power is used by an aircraft inverter or electric motor (e.g., information about how a controller will increase or decrease the power provided to the motor at a particular time in the future). The battery status information may include the state of charge, actual voltage, and / or current flowing in or out of the system's batteries or supercapacitors.

[0104] Thus, in operation 1504, the controller can determine how the power output of the hybrid generator should be adjusted to maintain the desired voltage range on the DC bus. For example, if the battery charge level is too low and in danger of not being able to maintain the desired voltage, in operation 1506, the controller can send instructions to increase the power output of the hybrid generator so that there is enough power to charge the battery. In another example, if the aircraft motors are using or are expected to need significantly more power than they are currently using, in operation 1506, the controller can send instructions to increase the power output of the hybrid generator. Similarly, the power output can also be decreased. In either case, the controller can adjust this overall power output to the DC bus by changing the RPM provided by the engine to the generator. Thus, while the battery pack and supercapacitor may reduce the need to adjust the power output of the hybrid generator set in real time, some control or adjustment of the RPM and therefore the output power to the DC bus may still be desirable in various embodiments, since the battery pack and / or supercapacitor can maintain the DC bus at a desired voltage level.

[0105] 16 is a flow chart illustrating an example method 1600 for maintaining a stable DC bus voltage based on measurements by a hybrid gen set-level controller, according to an example embodiment. Method 1600 is similar to method 1600, except that method 1600 contemplates measurements that may be performed by the hybrid gen set controller itself (e.g., controller 1462) rather than receiving such measurements or information from another controller (e.g., an aircraft system-wide controller, such as controller 1480 of FIG. 14).

[0106] In operation 1602, an aspect of the power available at or flowing through the DC bus is measured by the controller. In the event that the DC bus is measurable by the system-wide aircraft controller, operation 1602 may also be performed by the system-wide aircraft controller. Similarly, in the event that the battery and / or supercapacitor is packaged as part of the hybrid generator set rather than being located as part of the overall aircraft system, the controller may also measure the battery / supercapacitor status (e.g., state of charge, current, voltage, etc.) in operation 1602. In operation 304 (1604), the controller determines how to adjust the power output of the hybrid generator based on the measurements. For example, if the DC bus voltage is approaching outside of a desired range, it may be desirable in operation 306 (1606) to send instructions to components of the hybrid generator set to adjust the power output of the hybrid generator set based on the determination in operation 304 (1604) so ​​that the DC bus voltage remains within a desired voltage range.

[0107] Thrust control across multiple operating modes As described herein, an example of a hybrid electric power plant includes an engine, a motor / generator, a high voltage battery pack, a parallel hybrid output shaft operatively connected to a propeller, fan, or gearbox, and a high voltage connection that allows the engine's power output to be split or mixed between series electrical power generation and direct shaft power. As also described herein, such an architecture provides multiple different modes of operation. In some embodiments (e.g., as shown in FIG. 14), instead of a parallel hybrid output shaft being used to provide mechanical power to the propeller, fan, gearbox, etc. of the aircraft's propulsion system, the output shaft can power a generator (e.g., generator / motor 121, 185, 235, 1465 as described herein). Thus, in various embodiments, the thrust control described herein can have a usable range associated with such a system.

[0108] In aviation-based systems, it may be desirable to reduce the pilot's workload, whether mental or physical. Piloting an aircraft can require significant concentration, and a system that can provide reduced workload, reduced judgment, reduced need for memorization or use of checklists, etc., would reduce the likelihood of the pilot or operator making an error. In an automatic or manual lever system where a hybrid-electric powerplant controller receives a request for thrust from an overall aircraft system controller, the method as described herein also provides for a simpler design and / or interoperability with more aircraft, since the aircraft controller can simply request a given thrust level without having to program the aircraft controller to understand the various flight modes made available by the hybrid-electric powerplant.

[0109] The pilot may also find it advantageous to use thrust levers. Levers in an aircraft cockpit may include throttle levers used to direct the output power from one or more engines. The levers also relate to propeller thrust, with forward movement resulting in more thrust, faster climb, and / or faster cruise speed. Thus, in embodiments including physical levers, the pilot has the advantage that he or she is already familiar with the mechanisms for controlling the powerplant to increase thrust without having to retrain the pilot on the multiple operating modes of the hybrid-electric powerplant described herein.

[0110] Thus, the embodiments described herein also describe the physical layout of thrust levers across at least two operating modes of the series / parallel hybrid power plant, as well as aspects of the controller underlying this system. For example, as shown in FIG. 2A, controller 205 of flexible architecture 201 may receive a signal from main aircraft controller 220 indicating a request for a given thrust level, which may be related to the physical position of the lever or may be calculated by main aircraft controller 220 or another computing device.

[0111] In a first range (e.g., a first subset of positions in the full range of positions) (e.g., parallel hybrid generation mode), toward the lower end of the first range, the system can begin by supplying high voltage current (power) to the high voltage bus to drive the distributed electric propulsion. The engine RPM is at a high set point that provides maximum engine efficiency, and the motor / generators can be controlled to maintain the bus voltage. That is, the electrical output matches the aircraft load and the voltage is stable. The engine power can range from low power to maximum power and is determined only by the load on the HV bus. The first range can be shown, for example, as range 1705 in FIG. 17.

[0112] From this state, if the pilot wishes to engage the output shaft to rotate a mechanical device such as a thruster propeller or gearbox to drive the rotor, the pilot can begin to move the thrust lever forward, requesting that the direct drive shaft output thrust. If the electrical load does not already require full engine power, and unless the additional shaft power required also requires full engine power, moving the lever forward will initiate power blending while other automation present in the system maintains the bus voltage (and therefore DC output current). Power will still be required to maintain the HV bus, and power will also begin to flow to the output shaft, with the engine providing both mechanical shaft and electrical power simultaneously.

[0113] This continues until the point at which maximum power available from the engine is requested (e.g., point 1715 in FIG. 17) (both via the thrust levers and the deterministic automatic control and maintenance of bus voltage when DC current loads are present). At this thrust request, the parallel hybrid system output may be maximized either through wide open throttle, maximum machine fuel rack, or other such engine control range. This coincides with the midpoint of the thrust lever range labeled as the mode switch in FIG. 17.

[0114] If the DC current load from the distributed electric propulsion is not mitigated, additional thrust demands (increased shaft power to the propulsion propellers or gearbox) may require power to flow from the battery pack to the HV bus. At this stage (e.g., range 1710 in FIG. 17 ), designated auxiliary power mode, more engine power is provided to the output shaft and the power demands of the HV bus are met partly by the generator and partly by the battery pack.

[0115] This operation may continue until the pilot receives a warning regarding the performance and safety of the battery pack. Such warnings may relate to the state of charge (SOC), HV bus voltage (which decreases as the battery discharges), or battery temperature due to extended discharge. Once a limitation is detected and reported, the pilot or operator may select to reduce the thrust request from the output shaft. The pilot or operator may then lower the thrust lever position to rebalance the system and potentially return to a phase of automatically charging the battery pack from the hybrid power plant system (e.g., a first range called parallel hybrid generation mode). In various embodiments, the particular selection may also be made automatically by an on-board processor or controller. For example, if the HV bus voltage falls below a threshold, if the battery temperature is above a predefined threshold, if the state of charge (SOC) falls below a predefined threshold, etc., the processor or controller may automatically control which power output mode (or range of FIG. 17) the system is in, regardless of whether the physical controller operated by the pilot or other controller is within a particular range.

[0116] In various embodiments, additional or different operating modes of the hybrid-electric power plant may be incorporated into the operation in response to the lever and / or thrust level requests from the controller. For example, such embodiments may incorporate more than two different modes or may incorporate modes other than those shown in and described with respect to FIG. 17. For example, a third mode may be referred to as a whisper mode in which the engine is not operational and the motor / generator is powered by the battery pack to drive the mechanical output shaft. In such a mode, a lower overall power may be output than the two modes described above. Thus, such a mode may be applied at the lowest range of lever or thrust requests, with one or more other modes associated with other operating ranges of the lever or thrust requests.

[0117] For example, FIG. 18 illustrates example operating modes 1800 in which an example hybrid architecture can be controlled. While FIG. 17 illustrates two modes, FIG. 18 illustrates at least three modes, with a dashed line illustrating a possible fourth mode that can be implemented in an embodiment. The first threshold 1808, the second threshold 1810, and the third threshold 1812 can represent different levels of desired total power output of the system, which can be in the form of electrical or mechanical power output. For example, if the system includes the hybrid generator set 1461 of FIG. 14, the total power output can be the total amount of power provided to the bus 1467 by the combination of the generator 1465 and the battery packs 1482, 1484. When the desired amount of output power exceeds the first threshold 1808, the system transitions from the first operating mode 1802 to the second operating mode 1804. Similarly, as the desired amount of power moves from the second mode 1804 towards the third mode of operation 1806 (less power being delivered in areas of the second mode 1804 closest to the first mode 1802 and more power being delivered in areas of the second mode 1804 closer to the third mode 1806), at a second threshold 1810, the system may transition to the third mode of operation 1806. A similar effect occurs if the desired amount of power exceeds the third threshold and the system may transition to the fourth mode 1806.

[0118] In various embodiments, the modes may be associated with different outputs or modes, as described herein. For example, a first mode 1802 may be a mode that outputs power using only battery power. A second mode 1804 may be a mode in which power from the engine is output both to the bus (e.g., to charge the battery) and mechanically to the propulsion mechanism. In the example of FIG. 14, in such a second mode 1804, all power from the engine 1463 and generator 1465 is output to the bus 1467, some of the power is used by the electric motors 1474, 1478, some of which may be used to charge the batteries 1482, 1484. In a third mode, both power from the engine and power from the batteries are used to power the propulsion device (e.g., where the engine 1463 and generator 1465 and the batteries 1482, 1484 power the electric motors 1474, 1478).

[0119] Other modes described herein may be associated with any of the first, second, third, fourth, etc. modes of FIG. 14 in various embodiments. For example, the modes may include a hybrid generator mode as described herein, in which the engine may engage the power shaft using a clutch to drive the generator / motor and output power from the generator / motor. Another mode may be a direct drive engine mode as described herein, in which the engine drives the power shaft to mechanically or otherwise power the propulsion mechanism, while the power shaft rotates within the generator / motor and does not receive or output power at the generator / motor power input / output. Another mode may be an augmented thrust mode as described herein, in which both the engine and the generator / motor are used to simultaneously drive the power shaft and send power to the propulsion mechanism. Another mode may be a direct drive generator / motor mode as described herein, in which only the generator / motor may power the propulsion mechanism based on power received at the power input / output (e.g., from a battery pack). Another mode may be a split engine power mode, as described herein, in which the engine is used to drive a power shaft and a generator / motor to output power via a power input / output and a power shaft.

[0120] 19 is a schematic diagram of an example computing environment including a general-purpose computing system environment 100 such as a desktop computer, laptop, smartphone, tablet, or other device capable of executing instructions, such as those stored in a non-transitory computer-readable medium. Various computing devices disclosed herein (e.g., processor / controller 205, controller 220, processor / controller 280, hybrid generator set controller 1462, aircraft main controller 1480, or other computing devices in communication with controllers that may be part of other components of an aircraft) may be similar to computing system 100 or may include some components of computing system 100. Additionally, while described and illustrated in the context of a single computing system 100, those skilled in the art will also appreciate that various tasks described below may be performed in a distributed environment including multiple computing systems 100 linked via a local or wide area network, where executable instructions may be associated with and / or executed by one or more of the multiple computing systems 100.

[0121] In its most basic configuration, the computing system environment 100 typically includes at least one processing unit 102 and at least one memory 104, which may be linked via a bus 106. Depending on the exact configuration and type of computing system environment, the memory 104 may be volatile (such as RAM 110), non-volatile (such as ROM 108, flash memory, etc.), or a combination of the two. The computing system environment 100 may have additional features and / or functionality. For example, the computing system environment 100 may also include additional storage (erasable and non-erasable) including, but not limited to, magnetic or optical disks, tape drives, and / or flash drives. Such additional memory devices may be made accessible to the computing system environment 100, for example, via a hard disk drive interface 112, a magnetic disk drive interface 114, and / or an optical disk drive interface 116. As will be appreciated, these devices, each linked to the system bus 306, enable reading and writing to the hard disk 118, reading and writing to the erasable magnetic disk 120, and / or reading and writing to the erasable optical disk 122, such as a CD / DVD ROM or other optical medium. The drive interfaces and their associated computer readable media enable non-volatile storage of computer readable instructions, data structures, program modules, and other data for the computing system environment 100. Those skilled in the art will appreciate that other types of computer readable media capable of storing data may also be used for this same purpose. Examples of such media devices include, but are not limited to, magnetic cassettes, flash memory cards, digital video disks, Bernoulli cartridges, random access memory, nano drives, memory sticks, other read / write memory and / or read-only memory, and / or other methods or techniques for storing information, such as computer readable instructions, data structures, program modules, other data, and the like. Such computer storage media may be part of the computing system environment 100.

[0122] A number of program modules may be stored in one or more memory / media devices. For example, a basic input / output system (BIOS) 124, containing the basic routines that help to transfer information between elements within the computing system environment 100, such as during start-up, may be stored in ROM 108. Similarly, RAM 110, hard drive 118, and / or peripheral memory devices may be used to store computer-executable instructions including an operating system 126, one or more application programs 128 (which may include functionality disclosed herein, for example), other program modules 130, and / or program data 122. Additionally, computer-executable instructions may be downloaded to the computing environment 100 as needed, for example via a network connection.

[0123] An end user may enter commands and information into the computing system environment 100 through input devices such as a keyboard 134 and / or a pointing device 136. Although not shown, other input devices may include a microphone, joystick, game pad, scanner, etc. These and other input devices are typically connected to the processing unit 102 through a peripheral interface 138, which is coupled to the bus 106. The input devices may be connected directly or indirectly to the processor 102 through an interface such as a parallel port, game port, Firewire, or universal serial bus (USB). To display information from the computing system environment 100, a monitor 140 or other type of display device may also be connected to the bus 106 through an interface such as a video adapter 132. In addition to the monitor 140, the computing system environment 100 may also include other peripheral output devices (not shown), such as speakers and printers.

[0124] The computing system environment 100 may also utilize logical connections to one or more computing system environments. Communications between the computing system environment 100 and a remote computing system environment are exchanged through another processing device, such as a network router 152 responsible for network routing. Communications with the network router 152 may be performed through a network interface element 154. Thus, it will be understood that within such a network environment, such as the Internet, the World Wide Web, a LAN, or other similar types of wired or wireless networks, program modules depicted relative to the computing system environment 100, or portions thereof, may be stored in a memory storage device of the computing system environment 100.

[0125] The computing system environment 100 may also include location determination hardware 186 for determining the location of the computing system environment 100. In some cases, the location determination hardware 156 may include, for example, a GPS antenna, an RFID chip or reader, a WiFi antenna, or other computing hardware that can be used to capture or transmit signals that can be used to determine the location of the computing system environment 100.

[0126] Although specific embodiments have been described in this disclosure, it will be understood that the claims are not intended to be limited to these embodiments, unless expressly set forth in the claims. Rather, the disclosure is intended to cover alternatives, modifications, and equivalents that may be included within the spirit and scope of the disclosure. Furthermore, in the detailed description of the disclosure, numerous specific details are set forth in order to provide a thorough understanding of the disclosed embodiments. However, it will be apparent to those skilled in the art that systems and methods consistent with the present disclosure may be practiced without these specific details. In other instances, well-known methods, procedures, components, and circuits have not been described in detail so as not to unnecessarily obscure various aspects of the present disclosure.

[0127] Some of the detailed descriptions of this disclosure are presented in terms of procedures, logic blocks, processes, and other symbolic representations of operations on data bits within a computer or digital system memory. These descriptions and representations are the means used by those skilled in the data processing arts to most effectively convey the substance of their work to others skilled in the art. A procedure, logic block, process, or the like, is here, and generally, conceived to be a self-consistent sequence of steps or instructions leading to a desired result. The steps are those requiring physical manipulations of physical quantities. Usually, though not necessarily, these physical manipulations take the form of electrical or magnetic data capable of being stored, transferred, combined, compared, and otherwise manipulated in a computer system or similar electronic computing device. For convenience, and with reference to common usage, such data will be referred to as bits, values, elements, symbols, characters, terms, numbers, or the like with reference to the various embodiments presently disclosed.

[0128] However, it should be noted that these terms should be interpreted as referring to physical operations and quantities, are merely convenient labels, and should be further interpreted in light of the terms commonly used in the art. Unless otherwise indicated, as will be apparent from the discussion herein, throughout the discussion of the present embodiments, discussions using terms such as "determine" or "output" or "transmit" or "record" or "locate" or "store" or "display" or "receive" or "recognize" or "use" or "generate" or "provide" or "access" or "confirm" or "notify" or "deliver" are understood to refer to the operations and processes of a computer system or similar electronic computing device that manipulates and transforms data. Data is represented as physical quantities (electronic quantities) in the registers and memory of the computer system, and is transformed to other data that are also represented as physical quantities in the memory or registers of the computer system, or in other information storage, transmission, or display devices as described herein or understood by those skilled in the art.

[0129] In an exemplary embodiment, any of the operations described herein may be implemented, at least in part, as computer readable instructions stored on a computer readable medium or in a memory that, when executed by a processor, causes a computing device to perform an operation.

[0130] The foregoing description of the exemplary embodiments has been presented for purposes of illustration and description. It is not intended to be exhaustive or to be limited to the precise form disclosed, and modifications and variations may be possible in light of the above teachings or from practice of the disclosed embodiments. It is intended that the scope of the invention be defined by the claims appended hereto and their equivalents.

Claims

1. 1. A lever for adjusting the output of a hybrid electric powerplant of an aircraft, the lever comprising: a lever configured to move through a range of positions; Movement of the lever regulates the output of the hybrid-electric power plant between at least two operating modes; At a first subset of locations within the total range of locations, the hybrid-electric power plant: operating an engine having a mechanical output; outputting a first electrical energy from a motor / generator driven by the mechanical output of the engine; and a propulsion mechanism is driven by the mechanical output of the engine; At a second subset of locations within the total range of locations, the hybrid-electric power plant: operating the engine having the mechanical output; receiving second electrical energy at the motor / generator; driving the mechanical output with the motor / generator using the second electrical energy; and configured to drive the propulsion mechanism with the mechanical output; lever.

2. 10. A method for adjusting the output of an aircraft hybrid-electric powerplant using the lever of claim 1.

3. 10. A non-transitory computer-readable medium having stored thereon instructions that, when executed by a computing device, cause the computing device to perform operations for adjusting an output of a hybrid-electric powerplant of an aircraft using the lever of claim 1.

4. the first subset of locations represents a first contiguous group of locations; The lever of claim 1 , wherein the second subset of positions represents a second contiguous group of positions.

5. The lever of claim 4 , wherein one of the first subset of positions is adjacent to one of the second subset of positions.

6. The lever of claim 1 , wherein movement of the lever regulates the output of the hybrid-electric power plant among three or more operating modes.

7. 1. A thrust control system for regulating an output of a hybrid electric powerplant of an aircraft, the thrust control system comprising: a controller input configured to receive a command; the controller is configured to set an operating mode of the hybrid system based on the command, the operating mode including an output mode of the hybrid-electric power plant, and there are at least two operating modes; Furthermore, upon receiving a first command at the input, the hybrid-electric power plant: operating an engine having a mechanical output; configured to output first electrical energy from a motor / generator driven by the mechanical output of the engine, the first electrical energy being output to an electric propulsion motor of the aircraft and an aircraft battery; Upon receiving a second command at the input, the hybrid-electric power plant: configured to output second electrical energy from the motor / generator, the second electrical energy being output to the electric propulsion motor of the aircraft rather than to the battery of the aircraft; Thrust control system.

8. The input includes a lever configured to move in response to a force from a pilot or operator, with different positions used as commands corresponding to different modes; or 8. The thrust control system of claim 7, wherein the input includes an electrical connection to a computerized flight control system, and the controller is configured to receive electronic commands from the computerized flight control system corresponding to different modes.

9. 8. The thrust control system of claim 7, wherein upon receiving the first command, the hybrid-electric powerplant operates in a first of the at least two modes of operation, and upon receiving the second command, the hybrid-electric powerplant operates in a second of the at least two modes of operation.

10. In the second mode, the battery is configured to output a third electrical energy to the electric propulsion motors of the aircraft; or 10. The thrust control system of claim 9, wherein the motor / generator is driven by the mechanical output of the motor / generator.

11. The thrust control system of claim 7 , wherein the electric traction motor is connected to an inverter, the inverter being connected to a direct current (DC) bus.

12. the battery is connected to the DC bus; the inverter is a first inverter, and the motor / generator is connected to a second inverter; The thrust control system of claim 11 , wherein the second inverter is connected to the DC bus.

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