Systems and methods for aircraft flight control

Abstract

Aspects of the present disclosure generally relate to systems and methods for flight control of aircraft driven by electric propulsion systems and other types of vehicles. In some embodiments, an aircraft flight control system is disclosed, configured to determine one or more desired commands for an electric aircraft; determine at least one reference command based on the desired commands and one or more aircraft conditions; monitor the energy state of a plurality of battery packs, wherein at least one first battery pack of the plurality of battery packs is electrically isolated from at least two battery packs of the plurality of battery packs; adjust at least one reference command based on the monitored energy state of the plurality of battery packs; generate control commands for a plurality of effectors based on the adjusted at least one reference command; and control the plurality of effectors according to the generated control commands to satisfy one or more desired commands of the electric aircraft.

Description

Technical Field

[0001] Cross - reference to Related Applications This disclosure claims priority to U.S. Provisional Application No. 63 / 512,784, filed on July 10, 2023, entitled "SYSTEMS AND METHOD FOR FLIGHT CONTROL OF EVTOL AIRCRAFT" (Attorney Docket No. 16499.6006 - 00000), the content of which is hereby incorporated in its entirety for all purposes.

[0002] This disclosure generally relates to powered aerial vehicles. More specifically, and without limitation, this disclosure relates to technological innovations in aircraft driven by electric propulsion systems. Certain aspects of this disclosure generally relate to systems and methods for flight control of aircraft driven by electric propulsion systems and for flight control in other types of vehicles, as well as to flight control of aircraft in flight simulators and video games.

Background Art

[0003] The inventors of the present invention have come to recognize several problems that may be associated with the flight control of aircraft, including tiltrotor aircraft that use electric propulsion systems or hybrid electric propulsion systems (hereinafter referred to as electric propulsion units or "EPUs"). Vertical take-off and landing (VTOL) aircraft are aircraft that can take off and land vertically and hover, providing the ability to carry passengers directly to their destinations. Helicopters are VTOL aircraft that generate lift entirely through their rotors. Some VTOL aircraft have wings and a propulsion system, the propulsion system allowing the wings to provide the necessary lift during forward flight. Some airfoil VTOL aircraft use separate thrust systems for vertical thrust used during take-off and landing and forward thrust used during cruising. Other airfoil VTOL aircraft use a tiltable propulsion system that tilts between a vertical thrust position and a forward thrust position. Electric VTOL aircraft use electric propulsion units to provide thrust for vertical and forward flight. Many electric VTOL aircraft include a movable electric propulsion unit that can change the thrust vector of the propulsion unit, such as from an upward direction for vertical lift to a forward direction for forward flight. Many electric VTOL aircraft are overacted in that they have more actuator degrees of freedom than kinetic degrees of freedom. Control allocation is the problem of distributing control efforts among multiple actuators in an overacted system, such as optimizing energy and managing the temperatures of different aircraft components. Electric VTOL aircraft often contain more propulsion units and other actuators than conventional aircraft, and propulsion units and other actuators strongly influence multiple control axes. Therefore, electric VTOL aircraft can present greater control allocation problems than conventional aircraft. [Overview of the Initiative]

[0004] This disclosure generally relates to flight control of electric aircraft and other powered aerial vehicles. More specifically, and not limited to, this disclosure relates to technological innovations for tiltrotor aircraft using electric propulsion systems. Certain aspects of this disclosure, for example, relate to energy optimization and / or thermal management for aircraft.

[0005] One aspect of the present disclosure includes a method which comprises determining one or more desired directives for an aircraft; determining at least one reference directive based on the one or more desired directives and one or more aircraft conditions; monitoring the energy states of a plurality of battery packs of an electric aircraft, wherein at least one first battery pack of the plurality is electrically isolated from at least two battery packs of the plurality; adjusting at least one reference directive based on the monitored energy states of the plurality of battery packs; generating control directives for a plurality of effectors of an electric aircraft based on the adjusted at least one reference directive; and controlling the plurality of effectors in accordance with the generated control directives to satisfy one or more desired directives of the electric aircraft.

[0006] Another aspect of the present disclosure includes a system comprising at least one processor configured to determine at least one reference instruction based on one or more desired instructions and one or more aircraft conditions; monitor the energy state of a plurality of battery packs of an electric aircraft, wherein at least one first battery pack of the plurality is electrically isolated from at least one second battery pack of the plurality; adjust at least one reference instruction based on the monitored energy state of the plurality of battery packs; generate control instructions for a plurality of effectors of an electric aircraft based on the adjusted at least one reference instruction; and control the plurality of effectors according to the generated control instructions to satisfy one or more desired instructions of the electric aircraft.

[0007] Another aspect of the present disclosure includes a non-temporary computer-readable medium storing one or more instructions, which, when executed by at least one processor, cause at least one processor to perform an operation, the operation comprising determining one or more desired instructions for an electric aircraft; determining at least one reference instruction based on the one or more desired instructions and one or more aircraft conditions; monitoring the energy states of a plurality of battery packs of the electric aircraft, wherein at least one first battery pack of the plurality is electrically isolated from at least two battery packs of the plurality; adjusting at least one reference instruction based on the monitored energy states of the plurality of battery packs; generating control instructions for a plurality of effectors of the electric aircraft based on the adjusted at least one reference instruction; and controlling the plurality of effectors according to the generated control instructions to satisfy one or more desired instructions of the electric aircraft.

[0008] Another aspect of the present disclosure includes a method which includes determining one or more desired instructions for an electric aircraft; receiving engine information relating to at least one EPU among a plurality of EPUs of the electric aircraft, wherein the engine information includes at least one temperature associated with the at least one EPU; generating control instructions for a plurality of effectors of the electric aircraft based on the received engine information; and controlling the plurality of effectors in accordance with the generated control instructions to satisfy one or more desired instructions of the electric aircraft.

[0009] Another aspect of the present disclosure includes a system comprising at least one processor configured to determine one or more desired commands for an electric aircraft, receive engine information including at least one temperature associated with at least one of the EPUs among a plurality of EPUs of the electric aircraft, generate control commands for a plurality of effectors of the electric aircraft based on the received engine information, and control the plurality of effectors according to the generated control commands to satisfy one or more desired commands for the electric aircraft.

[0010] Another aspect of the present disclosure includes a non-temporary computer-readable medium storing one or more instructions, the one or more instructions causing at least one processor to perform an operation when executed by at least one processor, the operation including determining one or more desired instructions for an electric aircraft; receiving engine information relating to at least one of a plurality of EPUs of the electric aircraft, wherein the engine information includes at least one temperature associated with the at least one EPU; generating control instructions for a plurality of effectors of the electric aircraft based on the received engine information; and controlling the plurality of effectors in accordance with the generated control instructions to satisfy one or more desired instructions of the electric aircraft.

[0011] Another aspect of the present disclosure includes an aircraft engine, the engine including at least one processor configured to determine one or more temperatures associated with one or more components of the engine, and based on the determined one or more temperatures, estimate the remaining time corresponding to a prediction of when the engine will reach one or more predetermined limits for an electric engine at the current power setting.

[0012] Another aspect of the present disclosure includes a method which includes determining one or more desired instructions for an electric aircraft; receiving engine information relating to at least one EPU; receiving battery information relating to at least one battery pack; generating control instructions for a plurality of effectors of the electric aircraft based on the received engine information and battery information; and controlling the plurality of effectors in accordance with the generated control instructions to satisfy one or more desired instructions for the electric aircraft. [Brief explanation of the drawing]

[0013] [Figure 1] An exemplary VTOL aircraft consistent with the disclosed embodiments is shown.

[0014] [Figure 2] An exemplary VTOL aircraft consistent with the disclosed embodiments is shown.

[0015] [Figure 3] An exemplary top view of a VTOL aircraft, consistent with the disclosed embodiments, is shown.

[0016] [Figure 4] An exemplary propeller rotation of a VTOL aircraft, consistent with the disclosed embodiments, is illustrated.

[0017] [Figure 5] This shows an exemplary power connection in a VTOL aircraft, consistent with the disclosed embodiments.

[0018] [Figure 6] This is a schematic block diagram of an exemplary architecture and design of an electric propulsion and control system consistent with the disclosed embodiments.

[0019] [Figure 7] An exemplary top view of a VTOL aircraft, consistent with the disclosed embodiments, is shown.

[0020] [Figure 8] Schematic illustration of a flight control signaling architecture for controlling a control surface and associated actuators consistent with the disclosed embodiments.

[0021] [Figure 9A] Illustrative top view of a VTOL aircraft consistent with the disclosed embodiments. [Figure 9B] Illustrative top view of a VTOL aircraft consistent with the disclosed embodiments. [Figure 9C] Illustrative top view of a VTOL aircraft consistent with the disclosed embodiments. [Figure 9D] Illustrative top view of a VTOL aircraft consistent with the disclosed embodiments. [Figure 9E] Illustrative top view of a VTOL aircraft consistent with the disclosed embodiments.

[0022] [Figure 10] Functional block diagram of an exemplary control system of an electric VTOL aircraft consistent with the disclosed embodiments.

[0023] [Figure 11] Functional block diagram of an exemplary energy optimization method according to some embodiments.

[0024] [Figure 12] Exemplary method for energy optimization according to some embodiments.

[0025] [Figure 13A] Exemplary battery pack fault scenarios according to some embodiments. [Figure 13B] Exemplary battery pack fault scenarios according to some embodiments. [Figure 13C]This document illustrates exemplary battery pack failure scenarios in several embodiments. [Figure 13D] This document illustrates exemplary battery pack failure scenarios in several embodiments.

[0026] [Figure 14] This document presents exemplary scenarios for energy optimization through several embodiments. [Figure 15] This document presents exemplary scenarios for energy optimization through several embodiments. [Figure 16] This document presents exemplary scenarios for energy optimization through several embodiments. [Figure 17] This document presents exemplary scenarios for energy optimization through several embodiments. [Figure 18] This document presents exemplary scenarios for energy optimization through several embodiments.

[0027] [Figure 19] Exemplary temperature-versus-time plots are shown for several embodiments.

[0028] [Figure 20] The following are exemplary gauges related to engine temperature according to several embodiments.

[0029] [Figure 21] The following are exemplary temperature thresholds related to engine temperature in several embodiments.

[0030] [Figure 22] Block diagrams of exemplary aircraft control systems according to several embodiments are shown.

[0031] [Figure 23] The following describes exemplary methods for engine temperature control according to several embodiments.

[0032] [Figure 24] Block diagrams of exemplary control systems, including energy optimization and engine temperature optimization functions, according to several embodiments are shown.

[0033] [Figure 25A] This document presents exemplary scenarios for implementing thermal management and energy optimization through several embodiments. [Figure 25B] This document presents exemplary scenarios for implementing thermal management and energy optimization through several embodiments. [Modes for carrying out the invention]

[0034] This disclosure primarily describes systems, components, and technologies for use in aircraft. Aircraft may be piloted aircraft, unpiloted aircraft (e.g., UAVs), drones, helicopters, and / or airplanes. An aircraft includes a physical body and one or more components configured to enable the aircraft to fly (e.g., wings, tails, propellers, actuators, engines or motors, propulsion units, fuselages, effectors). An aircraft may include any configuration including at least one propeller. In some embodiments, an aircraft is driven (e.g., provided with thrust) by one or more electric propulsion systems (hereinafter referred to as electric propulsion units or "EPUs") which may include at least one engine, at least one rotor, at least one propeller, or any combination thereof. An aircraft may be fully electric, hybrid, or gas-powered. For example, in some embodiments, the aircraft is a tiltrotor configured for frequent (e.g., more than 50 flights per working day), short-duration flights (e.g., less than 100 miles per flight) over, into, and outside densely populated areas. The aircraft may be configured to carry 4 to 6 passengers or commuters who expect a comfortable experience with low noise and low vibration. Therefore, it is desirable to control the aircraft components in a manner that optimizes aircraft resources to improve aircraft performance (e.g., improving safety, energy efficiency, ride comfort, payload capacity, or structural integrity).

[0035] The disclosed embodiments provide novel and improved aircraft component configurations and / or identified design criteria that differ from those of conventional aircraft components, which are not found in conventional aircraft. By combining such alternative configurations and design criteria that address the shortcomings and challenges associated with conventional components, the embodiments disclosed herein relating to the configuration and design of various components for aircraft driven by electric propulsion systems (e.g., electric aircraft or hybrid electric aircraft) are obtained.

[0036] In some embodiments, an aircraft powered by the electric propulsion system of the present disclosure may be designed to be capable of both vertical and conventional takeoff and landing, with a distributed electric propulsion system that enables vertical, horizontal, and lateral flight, as well as transitions (e.g., transitions between vertical and horizontal flight). The aircraft may generate thrust by supplying high-voltage power to multiple engines of the distributed electric propulsion system, and the multiple engines may include components for converting high-voltage (HV) power into mechanical shaft power for rotating propellers.

[0037] Embodiments may include an electric engine (e.g., a motor) connected to an on-board power supply, which may include a device capable of storing energy, such as a battery or capacitor, and optionally include one or more systems for powering or generating electricity, such as a fuel-driven generator or a solar panel array. In some embodiments, the aircraft may include a hybrid aircraft that uses at least one of an electric-based energy source or a fuel-based energy source to power a distributed propulsion system. In some embodiments, the aircraft may be powered by one or more batteries, internal combustion engines (ICE), generators, turbine engines, or ducted fans.

[0038] The engines may be mounted directly to the wing or to one or more booms attached to the wing. The amount of thrust generated by each engine may be controlled by torque commands from the flight control system (FCS) via a digital communication interface to each engine. Embodiments may include forward-facing engines (and associated propellers) that can change their orientation or inclination.

[0039] The engine can also rotate the propeller clockwise or counterclockwise. In some embodiments, the difference in propeller rotation direction can be achieved using the engine rotation direction. In other embodiments, the engines may all rotate in the same direction, and gearing may be used to achieve different propeller rotation directions.

[0040] In some embodiments, an aircraft may have a number of engines in various combinations of forward-engine and rear-engine configurations. Forward-engine engines may be thought of as engines primarily located near the leading edge of the wing. Rear-engine engines may be thought of as engines primarily located near the trailing edge of the wing. For example, an aircraft may have any other combination of forward-engine and rear-engine configurations, including six forward-engine and six rear-engine configurations, five forward-engine and five rear-engine configurations, four forward-engine and four rear-engine configurations, three forward-engine and three rear-engine configurations, two forward-engine and two rear-engine configurations, or embodiments in which the number of forward-engine and rear-engine configurations are not equal.

[0041] In some embodiments, for vertical takeoff and landing (VTOL) missions, the forward and aft engines may provide vertical thrust during takeoff and landing. During the forward flight phase, the forward engine may provide horizontal thrust, while the aft engine's propeller may be retracted to a fixed position to minimize drag. The aft engine may be actively retracted while maintaining positional monitoring.

[0042] Transitions from vertical to horizontal flight and vice versa can be achieved via a tilt propeller subsystem. The tilt propeller subsystem can change the direction of thrust between a primarily vertical direction during the vertical flight phase (e.g., the hovering phase) and a horizontal or nearly horizontal direction during the forward flight cruising phase, based on the tilt of one or more propellers (e.g., determining the orientation of one or more propellers). A variable pitch mechanism can change the collective angle of the blades of the forward engine's propeller hub assembly for operation during flight phases such as the hovering phase, transition phase, and cruising phase. Vertical lift can be primarily vertical thrust (e.g., during the hovering phase). Horizontal thrust can be primarily horizontal thrust (e.g., during the cruising phase). In some embodiments, a flight phase (e.g., hovering, cruising, forward flight, takeoff, landing, transition to or from forward flight) may be defined by a combination of flight conditions (e.g., a combination of flight conditions within a specific range) which may include one or more of the following: airspeed, altitude, pitch angle (e.g., of the aircraft), tilt angle (e.g., of one or more propellers), roll angle, rotational speed (e.g., of the propellers), torque value, pilot commands, sensor measurements (e.g., accelerometer measurements), or any other values ​​indicating the current (e.g., experienced) or requested (e.g., commanded) state of at least some part of the aircraft.

[0043] In some embodiments, in conventional take-off and landing (CTOL) missions, the forward engines may provide horizontal thrust for fixed-wing take-off, cruising, and landing, while the wings may provide vertical lift. In some embodiments, the rear engines may not be used to generate thrust during CTOL missions, and the rear propellers may be retracted into a fixed position. In other embodiments, the rear engines may be used at reduced power to shorten the length of CTOL take-off or landing.

[0044] As detailed above, an aircraft embodiment may include many movable structural flight elements that enable the pilot to safely control the aircraft. Flight control surfaces (e.g., flaperons, flaps, ailerons, elevators, rudders, etc.) are crucial for controlling the aircraft's positioning. Changes in the orientation of these surfaces alter the airflow and pressure distribution around the aircraft, allowing the pilot to control the aircraft's motion on three axes of rotation. Similarly, control of propeller rotation and orientation can provide lift support (e.g., the lift required for vertical takeoff, landing, and hovering) and forward thrust necessary to move the aircraft through the air. The operation of each of these flight elements is important for the aircraft's safety and stability.

[0045] Herein, exemplary embodiments are given in detail, and examples are illustrated in the accompanying drawings. Where the following description is based on the drawings, unless otherwise specified, the same reference numerals in different drawings represent the same or similar elements. The implementations described in the following exemplary embodiments do not represent all implementations consistent with the present disclosure. Rather, they are merely examples of apparatus and methods consistent with embodiments relating to the subject matter described in the accompanying claims.

[0046] Figure 1 is an illustrative perspective view of an exemplary VTOL aircraft consistent with the disclosed embodiments. Figure 2 is an illustrative perspective view of an exemplary VTOL aircraft in an alternative configuration consistent with embodiments of the present disclosure. Figures 1 and 2 illustrate VTOL aircraft 100, 200 in cruising configuration and vertical takeoff, landing, and hovering configuration (also referred to herein as “lift” configuration), respectively, consistent with embodiments of the present disclosure. Elements corresponding to Figures 1 and 2 may have similar figures and may refer to similar elements of aircraft 100, 200. Aircraft 100, 200 may include fuselages 102, 202, wings 104, 204 mounted on fuselages 102, 202, and one or more rear stabilizers 106, 206 mounted on the rear of fuselages 102, 202. Multiple lift propellers 112, 212 may be mounted on wings 104, 204 and may be configured to provide lift for vertical takeoff, landing, and hovering. Multiple tilt propellers 114, 214 may be mounted on wings 104, 204 and may be tiltable (for example, configured to tilt or change direction) between a lift configuration in which these tilt propellers provide a portion of the lift required for vertical takeoff, landing, and hovering, as shown in Figure 2, and a cruising configuration in which these tilt propellers provide forward thrust to the aircraft 100 for horizontal flight, as shown in Figure 1. As used herein, the lift configuration of a tilt propeller refers to any tilt propeller orientation in which the tilt propeller thrust primarily provides lift to the aircraft, and the cruising configuration of a tilt propeller refers to any tilt propeller orientation in which the tilt propeller thrust primarily provides forward thrust to the aircraft.

[0047] In some embodiments, the lift propellers 112, 212 may be configured to provide only lift, with all horizontal thrust provided by the tilt propellers. For example, the lift propellers 112, 212 may be configured in a fixed position and may generate thrust only during the takeoff, landing, and hovering phases of flight. On the other hand, the tilt propellers 114, 214 may be tilted upward in a lift configuration in which the thrust from the propellers 114, 214 is directed downward to provide additional lift.

[0048] For forward flight, the tilt propellers 114, 214 can be tilted from their lift configuration to their cruising configuration. In other words, the orientation of the tilt propellers 114, 214 can vary from an orientation in which the thrust of the tilt propellers is directed downward (to provide lift during vertical takeoff, landing, and hovering) to an orientation in which the thrust of the tilt propellers is directed aft (to provide forward thrust to the aircraft 100, 200). The tilt propeller assembly for a particular electric engine can be tilted around an axis of rotation defined by the mounting point connecting the boom and the electric engine. When the aircraft 100, 200 is in full forward flight, lift can be fully provided by the wings 104, 204. On the other hand, in the cruising configuration, the lift propellers 112, 212 can be shut off. The blades 120, 220 of the lift propellers 112, 212 can be held in a low-drag position for aircraft cruising. In some embodiments, the lift propellers 112, 212 may each have two blades 120, 220 that can be locked in a minimum drag position where one blade is directly in front of the other blade while the aircraft is cruising, for example as illustrated in Figure 1. In some embodiments, the lift propellers 112, 212 may have three or more blades. In some embodiments, the tilt propellers 114, 214 may include more blades 116, 216 than the lift propellers 112, 212. For example, as illustrated in Figures 1 and 2, the lift propellers 112, 212 may each include, for example, two blades, while the tilt propellers 114, 214 may each include more blades, such as the five blades shown. In some embodiments, each of the tilt propellers 114, 214 may have two to five blades, and possibly more, depending on the design considerations and requirements of the aircraft.

[0049] In some embodiments, the aircraft may include a single wing 104, 204 on each side of the fuselage 102, 202 (or a single wing extending over the entire aircraft). At least a portion of the lift propellers 112, 212 may be located behind the wings 104, 204 (e.g., the propeller rotation point is behind the wing from a bird's-eye view), and at least a portion of the tilt propellers 114, 214 may be located in front of the wings 104, 204 (e.g., the propeller rotation point is in front of the wing from a bird's-eye view). In some embodiments, all of the lift propellers 112, 212 may be located behind the wings 104, 204, and all of the tilt propellers 114, 214 may be located in front of the wings 104, 204. According to some embodiments, all lift propellers 112, 212 and tilt propellers 114, 214 may be mounted on the wings, i.e., the lift propellers or tilt propellers may not be mounted on the fuselage. In some embodiments, all lift propellers 112, 212 may be located behind the wings 104, 204, and all tilt propellers 114, 214 may be located in front of the wings 104, 204. According to some embodiments, all lift propellers 112, 212 and tilt propellers 114, 214 may be located inside the edges of the wings 104, 204.

[0050] In some embodiments, lift propellers 112, 212 and tilt propellers 114, 214 may be mounted on the wings 104, 204 by booms 122, 222. Booms 122, 222 may be mounted below, above, and / or incorporated into the wing profile. In some embodiments, lift propellers 112, 212 and tilt propellers 114, 214 may be mounted directly on the wings 104, 204. In some embodiments, one lift propeller 112, 212 and one tilt propeller 114, 214 may be mounted on each boom 122, 222. Lift propellers 112, 212 may be mounted at the rear end of booms 122, 222, and tilt propellers 114, 214 may be mounted at the front end of booms 122, 222. In some embodiments, lift propellers 112, 212 may be mounted in fixed positions on booms 122, 222. In some embodiments, tilt propellers 114, 214 may be mounted to the front end of booms 122, 222 via hinges. The tilt propellers 114, 214 may be mounted on booms 122, 222 such that when in their cruising configuration, the tilt propellers 114, 214 are aligned with the body of booms 122, 222, forming a continuous extension of the front end of booms 122, 222 that minimizes drag for forward flight.

[0051] In some embodiments, the aircraft 100, 200 may include, for example, one wing on each side of the fuselage 102, 202, or a single wing extending across the entire aircraft. According to some embodiments, at least one wing 104, 204 is a high wing mounted on the upper side of the fuselage 102, 202. According to some embodiments, the wing includes control surfaces such as flaps, ailerons, and / or flaperons (for example, configured to perform both flap and aileron functions). According to some embodiments, the wings 104, 204 may have a profile that reduces drag during forward flight. In some embodiments, the wingtip profile may be curved and / or tapered to minimize drag.

[0052] In some embodiments, the rear stabilizers 106, 206 include control surfaces such as one or more rudders, one or more elevators, and / or one or more combined rudder-elevator configurations. The wing(s) may have any preferred design to provide lift, directionality, stability, and / or any other properties beneficial to the aircraft. In some embodiments, the wing has a tapered leading edge.

[0053] In some embodiments, the lift propellers 112, 212 or tilt propellers 114, 214 may be tilted relative to at least one other lift propeller 112, 212 or tilt propeller 114, 214, where tilt refers to the relative orientation of the axis of rotation of the lift propeller / tilt propeller around a line parallel to the longitudinal direction, similar to the roll degrees of freedom of an aircraft.

[0054] In some embodiments, one or more lift propellers 112, 212 and / or tilt propellers 114, 214 may be tilted relative to the aircraft cabin such that the axis of rotation of the propellers in the lift configuration is angled away from an axis perpendicular to the upper surface of the aircraft. For example, in some embodiments, the aircraft is a flying wing aircraft as shown in Figure 9E below, and some or all of the propellers are tilted away from the cabin.

[0055] Figure 3 is an illustrative top view of an exemplary VTOL aircraft consistent with embodiments of the present disclosure. The aircraft 300 shown in the figure may be a top view of aircraft 100 and 200 shown in Figures 1 and 2, respectively. As considered herein, the aircraft 300 may include 12 electric propulsion systems distributed across the aircraft 300. In some embodiments, the distribution of electric propulsion systems may include six forward electric propulsion systems 314 and six rear electric propulsion systems 312 mounted on forward and rear booms of the main wing 304 of the aircraft 300. In some embodiments, the forward electric propulsion systems may be mounted on the wing 304 by boom 322. In some embodiments, the rear electric propulsion systems may be mounted on the wing 304 by boom 324. In some embodiments, the length of the rear end of boom 324 from the wing 304 to the lift propeller (part of the electric propulsion system 312) may include similar rear ends of the boom 324 across a number of rear ends of the boom. In some embodiments, the length of the rear ends of the boom may vary, for example, across six rear ends. Furthermore, Figure 3 illustrates an exemplary embodiment of a VTOL aircraft 300, in which the forward propeller (part of the electric propulsion system 314) is oriented horizontally for horizontal flight and the rear propeller blades 320 are in a retracted position for the forward flight phase.

[0056] Figure 4 is a schematic diagram illustrating exemplary propeller rotation of a VTOL aircraft consistent with the disclosed embodiments. The aircraft 400 shown in the figure may be a top view of aircraft 100, 200, and 300 shown in Figures 1, 2, and 3, respectively. The aircraft 400 may include six forward electric propulsion systems, three of which are clockwise (CW) type 424 and the remaining three forward electric propulsion systems are counterclockwise (CCW) type 426. In some embodiments, three rearward electric propulsion systems may be counterclockwise (CCW) type 428 and the remaining three rearward electric propulsion systems may be CW type 430. Some embodiments may include an aircraft 400 having four forward electric propulsion systems and four rearward electric propulsion systems, each having two CW types and two CCW types. In some embodiments, the aircraft 400 may include a fuselage 402, wings 404 mounted on the fuselage 402, and one or more rear stabilizers 406 mounted on the rear of the fuselage 402. In some embodiments, each forward electric propulsion system may include a propeller blade 416. In some embodiments, each rear electric propulsion system may include a propeller blade 420. In some embodiments, the electric propulsion system may be mounted on the wings 404 by a boom 422. In some embodiments, propellers may be reversed relative to adjacent propellers to cancel torque steer generated by the rotation of the propellers and experienced by the aircraft's fuselage or wings. In some embodiments, the difference in rotation direction may be achieved using the engine rotation direction. In other embodiments, the engines may all rotate in the same direction, and gearing may be used to achieve different propeller rotation directions.

[0057] Some embodiments may include an aircraft 400 having a forward electric propulsion system and a rear electric propulsion system, where the quantities of CW type 424 and CCW type 426 are not equal between forward electric propulsion systems, between rear electric propulsion systems, or between forward electric propulsion systems and rear electric propulsion systems.

[0058] Figure 5 illustrates an exemplary eVTOL aircraft consistent with embodiments of the present disclosure. In some embodiments, the distributed electric propulsion system of the eVTOL aircraft 500 may include a high-voltage power supply (HVPS) system for supplying high-voltage (HV) power. The HVPS system may be configured to supply power on the aircraft 500 and distribute electrical energy stored in other systems on the aircraft 500, including an electric propulsion system (EPS) for converting the power into mechanical rotating shaft power to generate thrust (e.g., via a high-voltage channel, via a high-voltage bus). As shown in Figure 5, the HVPS system of the aircraft 500 may include six battery packs 520 (numbered B1 to B6 from left to right) installed in battery bays in the wings of the aircraft 500. In some embodiments, the six battery packs 520 may have the same design to simplify design, manufacturing, and logistics. The battery packs 520 may power one or more electric engines 510 (numbered E1 to E12).

[0059] In some embodiments, a single battery pack 520 can be electrically connected to and power multiple electric engines 510. For example, in some embodiments, the battery pack 520 can power electric engines 510 on either side of the longitudinal axis. In some embodiments, the battery pack 520 can power electric engines 510 on either side of the horizontal axis (for example, along one or more wings of an aircraft). In some embodiments, as shown in Figure 5, the battery pack 520 can power two diagonally opposed electric engines 510. For example, battery pack B1 can power electric engines E1 and E12. Battery pack B2 can power electric engines E5 and E8. Battery pack B3 can power electric engines E3 and E10. Battery pack B4 can power electric engines E4 and E9. Battery pack B5 can power electric engines E2 and E11. Battery pack B6 can power electric engines E6 and E7. Therefore, when the battery pack 520 is lost, the loss of lift is optimized, thereby reducing the impact on roll or pitch moment. In some embodiments, the battery pack 520 may power different configurations of the electric engine 510 to reduce the roll moment, pitch moment, or yaw moment that may be caused by the loss of the battery pack 520. For example, in some embodiments, the battery pack 520 may be connected to the electric engine 510 in any way that improves (e.g., optimizes) lift and / or thrust across the longitudinal and horizontal axes of the aircraft.

[0060] Furthermore, the HVPS system may include at least one crosslink 530 (e.g., a high-voltage bus) having a fuse that allows two or more battery packs 520 to be paired. Through the crosslink, power for the electric engine 510 can be shared among the paired battery packs 520. Thus, multiple battery packs 520 can power multiple electric engines 510 simultaneously. This arrangement provides redundancy and avoids a single point of failure, as each paired battery 520 can act as a backup for the other battery(s). In the event of a battery pack 520 failure, one or more connected battery packs 520 can continue to power the electric engine 510 connected to the failed battery pack.

[0061] In some embodiments, as shown in Figure 5, a pair of battery packs 520 may include two battery packs 520. In some embodiments, a pair of two battery packs 520 may power a total of four electric engines 510. For example, battery pack B1, which powers electric engines E1 and E12, may be cross-linked to battery pack B4, which powers electric engines E4 and E9 (for example, connected via a high-voltage bus). Battery pack B2, which powers electric engines E5 and E8, may be cross-linked to battery pack B5, which powers electric engines E2 and E11. Battery pack B3, which powers electric engines E3 and E10, may be cross-linked to battery pack B6, which powers electric engines E6 and E7.

[0062] A crosslink (e.g., a high-voltage bus) may connect two high-voltage channels, which may power one or more electric engines. In some embodiments, the crosslink may connect to the high-voltage channels of at least one (e.g., each) battery pack before the channels are split to power multiple electric engines (e.g., two electric engines).

[0063] In some embodiments, three or more battery packs 520 may be cross-linked together. For example, in some embodiments, three battery packs 520 may be cross-linked. Thus, in some embodiments, three battery packs 520 may power six electric engines 510. In some embodiments, four battery packs 520 may be cross-linked. Thus, in some embodiments, four battery packs 520 may power eight electric engines 510. In some embodiments, different arrangements of battery packs 520 and cross-linking may be selected to arbitrarily optimize the aircraft's power needs, system redundancy, and fault tolerance.

[0064] In some embodiments, each battery pack may include a battery management system housed within an HV distribution unit and a high-voltage junction box (HVJB). The battery management system may include one or more processors, microprocessors, and / or controllers. The BMS may be configured to monitor voltage, temperature, current, and isolation resistance. The BMS may be configured to protect against fault conditions (e.g., using fuses). As will be further detailed below, the BMS may communicate with various systems inside and outside the HVJB. The BMS may include a battery management unit (BMU) capable of receiving voltage, current, resistance, and temperature sensing signals from the cell stack assembly and / or the HV distribution unit.

[0065] The BMU can monitor output current for each connected load. The BMU can continuously monitor the battery status even when the battery is not installed in the aircraft 100. By monitoring the battery pack, the BMU can protect against conditions that adversely affect safety or performance, such as overvoltage, undervoltage, overheating, undertemperature, loss of electrical insulation, short circuits, and overcurrents. In addition, the BMU performs calculations of the battery pack's state of charge (SOC), state of health (SOH), fault conditions (e.g., short circuit or overcurrent), state of power (SOP), state of energy (SEO), and state of temperature (SOT). In some embodiments, the BMU may be configured to perform calculations of available energy (e.g., available energy) based on one or more of the SOC, SOH, SOP, SOE, or SOT. For example, the BMU may perform calculations of available energy for each battery pack, consistent with the disclosed embodiments. The BMU also controls and monitors bus precharging, provides fuse and contactor commands, and communicates with various systems inside and outside the HVJB. Furthermore, the BMU can communicate with aircraft switches and flight control systems and modify its operation based on received commands.

[0066] In some embodiments, the BMU may detect a fault event and transmit a command signal to blow a fuse. For example, the BMU may receive information about the conditions of a connected load at a point in the HVPS system (e.g., voltage, current, or temperature). Based on the received information, the BMU may determine a fault condition (e.g., because a value is outside a predetermined range) and transmit a command to blow the associated fuse. Thus, the fault condition can be disconnected from the rest of the HVPS circuit, protecting the remaining devices and wiring.

[0067] In some embodiments, battery packs may communicate with each other, for example, through a BMS. The battery packs may use information about the state of one or more paired battery packs within the battery pack unit to help determine whether an overcurrent condition has occurred. For example, the battery packs may determine an expected operating range (e.g., voltage, current, etc.) based on the state of the battery packs and the communicated state of the battery packs within the battery pack unit.

[0068] Figure 6 is a schematic block diagram of an exemplary architecture and design of the electric propulsion power and control system 600 consistent with the disclosed embodiments. The electric propulsion power and control system 600 includes an electric propulsion system 602 which may be configured to control an aircraft propeller. The electric propulsion system 602 may include an electric engine subsystem 604 which may supply torque to a propeller subsystem 606 via a shaft to generate thrust for the electric propulsion system 602. In some embodiments, the electric engine subsystem 604 may include receiving low-voltage direct current (LV DC) power from a low-voltage grid (LVS) 608. In some embodiments, the electric engine subsystem 604 may be configured to receive high-voltage (HV) power from a high-voltage power grid (HVPS) 610 which includes at least one battery or other device capable of storing energy. HV power may refer to power with a voltage lower than the voltage provided by the low-voltage grid (LVS) 608.

[0069] Some embodiments may include an electric propulsion system 602 which includes an electric engine subsystem 604 that receives signals from and transmits signals to the flight control system 612. In some embodiments, the flight control system (FCS) 612 may include a flight control computer (FCC) which can transmit commands to and receive status and data from the electric engine subsystem 604 using Controller Area Network ("CAN") data bus signals. The FCC may include devices configured to perform one or more operations for the aircraft (e.g., computation operations), such as at least one processor and memory components, consistent with the disclosed embodiments, wherein the at least one processor and memory component can store instructions that can be executed by at least one processor to perform the operations. While the CAN data bus signals are used between the flight control computer and the electric engine(s), it should be understood that some embodiments may include any form of communication that has the ability to send and receive data from the flight control computer to and from the electric engine(s). Some embodiments may include an electric engine subsystem 604 that can receive operating parameters, including speed, voltage, current, torque, temperature, vibration, propeller position, and / or any other values ​​of operating parameters, from the FCC in the FCS 612 and transmit the operating parameters to the FCC.

[0070] In some embodiments, the flight control system 612 may also include a tilt propeller system ("TPS") 614 capable of sending and receiving analog discrete data to and from the tilt propeller electric engine subsystem 604. The tilt propeller system (TPS) 614 may include a device that transmits operating parameters to the electric engine subsystem 604 and articulates the orientation of the propeller subsystem 606, thereby changing the direction of the tilt propeller thrust during various phases of flight using mechanical means such as a gearbox assembly, a linear actuator, and any other configuration of components for changing the orientation of the propeller subsystem 606. In some embodiments, the electric engine subsystem may transmit the orientation of the propeller system (e.g., the angle between lift and forward thrust) to the TPS 614 and / or FCS 612 (e.g., during flight).

[0071] In some embodiments, the flight control system may include a system capable of controlling a control surface and actuators associated with the control surface in an exemplary VTOL aircraft. Figure 7 illustrates a top view of an exemplary VTOL aircraft consistent with embodiments of the present disclosure. The aircraft 700 shown in the figure may be a top view of aircraft 100, 200 shown in Figures 1 and 2, respectively, in addition to the aircraft components described above with reference to Figure 3. In aircraft 700, the control surface may include a flaperon 712 and a rudder beta 714 in addition to the propeller blades considered above. The flaperon 712 may combine the functions of one or more flaps, one or more ailerons, and / or one or more spoilers. The rudder beta 714 may combine the functions of one or more rudders and / or one or more elevators. Additionally or alternatively, the control surface may include separate rudders and elevators. In aircraft 700, the actuators may include, in addition to the electric propulsion system discussed above, control surface actuators (CSAs) associated with flaperons 712 and rudder betas 714, as will be further discussed below with reference to Figure 8.

[0072] Figure 8 schematically illustrates a flight control signaling architecture 800 for controlling control surfaces and associated actuators, consistent with the disclosed embodiments. While Figure 7 illustrates 12 EPU inverters and associated propeller blades, 6 tilt propeller actuators (TPACs), 6 battery management systems (BMSs), 4 flaperons and associated control surface actuators (CSAs), and 6 rudder betas and associated CSAs, aircraft according to various embodiments may have any preferred number of these various elements. As shown in Figure 8, the control surfaces and actuators may be controlled by a combination of four flight control computers (FCCs): left FCC, lane A (L FCC-A) 801, left FCC, lane B (L FCC-B) 802, right FCC, lane A (R FCC-A) 803, and right FCC, lane A (R FCC-B) 804, although any other preferred number of FCCs may be available. Each FCC may control all control surfaces and actuators individually or in any combination thereof. In some embodiments, each FCC may include one or more hardware computing processors. In some embodiments, each FCC may utilize a single-threaded or multi-threaded computing process to perform the calculations required to control the control surfaces and actuators. In some embodiments, all computing processes required to control the control surfaces and actuators may be performed by a single flight control computer on a single computing thread.

[0073] The FCC may provide control signals to control surface actuators, including the EPU inverter 806, TPAC 808, BMS 809, flaperon CSA 810, and ladder beta CSA 811, via one or more bus systems. For different control surface actuators, the FCC may provide control signals such as voltage control signals or current control signals, and the control information may be encoded into binary, digital, or analog control signals. In some embodiments, each bus system may be a CAN bus system, e.g., left CAN bus 1, left CAN bus 2, right CAN bus 1, right CAN bus 2, center CAN bus 1, center CAN bus 2 (see Figure 8). In some embodiments, multiple FCCs may be configured to provide control signals via each CAN bus system, and each FCC may be configured to provide control signals via multiple CAN bus systems. In the exemplary architecture shown in Figure 8, for example, L FCC-A may provide control signals via left CAN bus 1 and right CAN bus 1, L FCC-B may provide control signals via left CAN bus 1 and central CAN bus 1, R FCC-A may provide control signals via central CAN bus 2 and right CAN bus 2, and R FCC-B may provide control signals via left CAN bus 2 and right CAN bus 2.

[0074] Figures 9A–9E illustrate exemplary top views of a VTOL aircraft consistent with embodiments of the present disclosure. There may be several design considerations (cost, weight, size, performance capabilities, etc.) that can influence the number and / or combination of tilt and lift propellers in a VTOL aircraft. As will be further discussed below, the number and orientation of aircraft components (e.g., effectors or actuators) can affect how energy is used and the temperatures of various aircraft components. Thus, a flight control system may control the aircraft in a manner that optimizes energy use and keeps various aircraft components at optimal temperatures by tuning the aircraft components in a particular manner (e.g., as considered in the disclosed embodiments).

[0075] Figure 9A illustrates an arrangement of the electric propulsion system 900 consistent with embodiments of the present disclosure. Referring to Figure 9A, the aircraft shown in the figure may be a top view of an exemplary aircraft (e.g., a VTOL aircraft). The aircraft may include 12 electric propulsion systems distributed across the aircraft. In some embodiments, the distribution of electric propulsion systems may include six forward electric propulsion systems 901, 902, 903, 904, 905, 906 and six rear electric propulsion systems 907, 908, 909, 910, 911, 912. In some embodiments, the six forward electric propulsion systems may be operably connected to tilt propellers, and the six rear electric propulsion systems may be operably connected to lift propellers. In other embodiments, the six forward electric propulsion systems and some rear electric propulsion systems may be operably connected to tilt propellers, and the remaining rear electric propulsion systems may be operably connected to lift propellers. In other embodiments, all forward and rear electric propulsion systems may be operably coupled to a tilt propeller.

[0076] Figure 9B illustrates an alternative arrangement of the electric propulsion system 940 consistent with embodiments of the present disclosure. Referring to Figure 9B, the aircraft shown in the figure may be a top view of an exemplary aircraft (e.g., a VTOL aircraft). The aircraft may include eight electric propulsion systems distributed across the aircraft. In some embodiments, the distribution of electric propulsion systems may include four forward electric propulsion systems 913, 914, 915, 916 and four rear electric propulsion systems 917, 918, 919, 920. In some embodiments, the four forward electric propulsion systems may be operably connected to tilt propellers, and the four rear electric propulsion systems may be operably connected to lift propellers. In other embodiments, the four forward electric propulsion systems and some rear electric propulsion systems may be operably connected to tilt propellers, and the remaining rear electric propulsion systems may be operably connected to lift propellers. In other embodiments, all forward and rear electric propulsion systems may be operably coupled to tilt propellers.

[0077] Figure 9C illustrates an alternative arrangement of the electric propulsion system 950 consistent with embodiments of the present disclosure. Referring to Figure 9C, the aircraft shown in the figure may be a top view of an exemplary aircraft (e.g., a VTOL aircraft). The aircraft may include six electric propulsion systems distributed across the aircraft. In some embodiments, the distribution of the electric propulsion systems may include a first set of four electric propulsion systems 921, 922, 923, and 924 coplanar in a first plane, and a second set of two electric propulsion systems 925 and 926 coplanar in a second plane. In some embodiments, the first set of electric propulsion systems 921-924 may be operably connected to tilt propellers, and the second set of electric propulsion systems 925 and 926 may be operably connected to lift propellers. In other embodiments, the first set of electric propulsion systems 921-924 and the second set of rear electric propulsion systems 925, 926 can all be operably connected to a tilt propeller.

[0078] Figure 9D illustrates an alternative arrangement of the electric propulsion system 960 consistent with embodiments of the present disclosure. Referring to Figure 9D, the aircraft shown in the figure may be a top view of an exemplary aircraft (e.g., a VTOL aircraft). The aircraft may include four electric propulsion systems spread out over the aircraft. In some embodiments, the distribution of electric propulsion systems may include four coplanar electric propulsion systems 927, 928, 929, and 930. In some embodiments, all of the electric propulsion systems may be operably connected to a tilt propeller.

[0079] Figure 9E illustrates an alternative arrangement of the electric propulsion system 970 consistent with embodiments of the present disclosure. Referring to Figure 9E, the aircraft shown in the figure may be a top view of an exemplary aircraft (e.g., a VTOL aircraft). The aircraft may include six electric propulsion systems spread out across the aircraft. For example, in some embodiments, the aircraft may include four forward electric propulsion systems 931, 932, 933, and 934 operably connected to a tilt propeller, and two rear electric propulsion systems 935 and 936 operably connected to a lift propeller. In some embodiments, the aircraft may include ten electric propulsion systems spread out across the aircraft. For example, in some embodiments, the aircraft may include six forward electric propulsion systems operably connected to a tilt propeller, and four rear electric propulsion systems operably connected to a lift propeller. In some embodiments, some or all of the rear electric propulsion systems may be operably connected to a tilt propeller.

[0080] As shown in Figure 9E, in some embodiments, the aircraft may have an all-wing configuration, such as a tailless fixed-wing aircraft without a distinct fuselage. In some embodiments, the aircraft may have an all-wing configuration in which the fuselage is integrated into the wings. In some embodiments, the tilt propeller may rotate in a plane above the aircraft body when the tilt propeller is operating in a lift configuration.

[0081] As disclosed herein, forward and rear electric propulsion systems may be of the clockwise (CW) or counterclockwise (CCW) type. Some embodiments may include a variety of forward electric propulsion systems having a mixture of both CW and CCW types. In some embodiments, the rear electric propulsion system may have a mixture of CW and CCW type systems among the rear electric propulsion systems. In some embodiments, each electric propulsion system may be fixed as either clockwise (CW) or counterclockwise (CCW) type, while in other embodiments, one or more electric propulsion systems may vary between clockwise (CW) and counterclockwise (CCW) rotation.

[0082] Figure 10 illustrates a functional block diagram of an exemplary aircraft control system 1000 consistent with the disclosed embodiments. System 1000 may be implemented by at least one processor (e.g., at least one microprocessor-based controller) configured to execute software code stored in a storage medium (e.g., computer-readable medium, non-temporary computer-readable medium) to implement the functions described herein. System 1000 may also be implemented in hardware, or a combination of hardware and software. System 1000 may be implemented as part of an aircraft flight control system (e.g., part of FCS612 in Figure 6) and may be configured to repeatedly perform a single step or sequence until a desired or commanded result is obtained. It should be understood that many conventional functions of control systems are not shown in Figure 10 for the sake of clarity. System 1000 further includes one or more storage media storing models, functions, tables, and / or arbitrary information for performing the disclosed processes. As further described below, any box or each box representing the command models 1004, 1006, 1008, 1010, feedback 1012, 1016, 1018, 1022, feedforward 1014, 1020, outer loop allocation 1024, 1026, inner loop control law 1028, control allocation 1029, and DCPS 1033 may represent or contain modules(multiple), scripts(multiple), functions(multiple), applications(multiple), and / or programs(multiple) executed by the processor(multiple) and / or microprocessor(multiple) of system 1000. The complexity and interoperability of the functional block diagram in Figure 10 is understood to be impossible, or at least impractical, to be effectively implemented by a human user, especially considering that these functionalities are implemented during aircraft flight (including takeoff or landing) (e.g., in real time).Unless otherwise specified herein, “desired” with respect to instructions for an aircraft, aircraft parameters, or other aircraft characteristics may refer to characteristics that are input (e.g., by a pilot input device, system module, etc.) as part of a process for controlling the computing-related operations of the aircraft, requested, and / or otherwise existing.

[0083] In some embodiments, the control system 1000 may be configured based on one or more flight control laws. A flight control law may include a set of algorithms, models, and / or rules configured to control the behavior of the aircraft (e.g., control or influence one or more effectors of the aircraft) in response to one or more pilot inputs and external factors. In some embodiments, a flight control law may be configured to achieve at least one of desired flight characteristics, stability, or performance. For example, a flight control law may be configured to ensure the stability and controllability of the aircraft by controlling how the aircraft responds to at least one of one or more pilot inputs, vehicle dynamics (e.g., turbulence, gusts, etc.), or changes in flight conditions (e.g., altitude, airspeed, angle of attack, torque of one or more propellers).

[0084] System 1000 may receive at least one pilot input and detect one or more inputs from pilot input devices configured to generate or influence signals. Pilot inputs may be generated by and / or received from aircraft input devices or mechanisms, such as buttons, switches, sticks, sliders, interceptors, or any other devices configured to generate or influence signals based on physical actions from the pilot. For example, pilot input devices may include one or more right interceptors (e.g., moving left / right 1002a and / or forward / backward 1002e), left interceptors (e.g., moving left / right 1002c and / or forward / backward 1002g), and / or left interceptor switches 1002f. In some embodiments, the pilot input device may include an interface with an autopilot system (e.g., display screens, switches, buttons, levers, and / or other interfaces). Optionally, system 1000 may further detect inputs from the autopilot system, such as autopilot roll commands 1002b, autopilot climb commands 1002d, and / or other commands for controlling the aircraft.

[0085] In some embodiments, one or more inputs may include at least one of the following: position and / or rate of the right inceptor and / or left inceptor; signals received from switches on the inceptor (e.g., response type change command, trim input, reference input, backup control input, etc.); measured aircraft status and environmental conditions based on data received from one or more sensors of the aircraft (e.g., measured load factor, airspeed, roll angle, pitch angle, actuator status, battery status, aerodynamic parameters, temperature, gusts, etc.); obstacles (e.g., presence or absence of other aircraft and / or debris); and aircraft mode (e.g., taxiing on the ground, takeoff, airborne). For example, right inceptor L / R 1002a may include the lateral position and / or rate of the right inceptor (e.g., an inceptor positioned to the right of another inceptor, and / or an inceptor positioned to the right of the pilot area), autopilot roll command 1002b may include the roll signal received in autopilot mode, left inceptor L / R 1002c may include the left inceptor (e.g., an inceptor positioned to the left of another inceptor, and / or an inceptor positioned to the left of the pilot area) The autopilot climb command 1002d may include the lateral position and / or rate of the positioned inceptor, the right inceptor F / A 1002e may include the longitudinal position and / or rate of the right inceptor, the left inceptor switch 1002f may include a signal from the switch to enable or disable the automatic transition function 1003, and the left inceptor F / A 1002g may include the longitudinal position and / or rate of the left inceptor.

[0086] At least one (e.g., each) input may contain data such as those listed above (e.g., signals from switches, measured aircraft status, aircraft mode, etc.). Actuator status may include actuator health indicators that indicate a degradation in actuator performance, which may limit the ability of a given actuator to satisfy an actuator command, and may include hardware limitations of the actuator, such as movement limits, speed limits, response time limits. Actuator status may be used to determine boundaries (e.g., minimum / maximum) for individual actuator commands. Battery status may correspond to the remaining energy of the aircraft's battery pack, which may be monitored when control allocation 1029 considers optimizing the battery pack's energy state. Aerodynamic parameters may be parameters derived from aerodynamic and acoustic modeling and may be based on the actuator's Jacobian matrix (e.g., describing how changes in the actuator (e.g., in response to actuator / effector commands) affect the aircraft's overall motion) and actuator status. At least one (e.g., each) input received from the interceptor may indicate the aircraft's heading or corresponding adjustment of output power.

[0087] Command models 1004, 1006, 1008, and 1010 may be configured to determine the shape of an ideal aircraft response (e.g., aggression, slew rate, damping, overshoot, etc.). For example, at least one of the command models 1004, 1006, 1008, and 1010 (e.g., each command model) may be configured to receive and interpret at least one of the inputs 1002a, 1002b, 1002c, 1002d, 1002e, 1002f, and 1002g, and in response to it, calculate the corresponding changes in the aircraft orientation, heading, and thrust, or combinations thereof, using an integrator (not shown). In some embodiments, the right inceptor L / R 1002a and autopilot roll command 1002b may be supplied to the turn rate command model 1004, the left inceptor L / R 1002c may be supplied to the lateral speed command model 1006, the autopilot climb command 1002d and right inceptor F / A 1002e may be supplied to the climb command model 1008, and the left inceptor F / A 1002g may be supplied to the forward speed command model 1010. In some embodiments, the output from the automatic transition function 1003 may be supplied to at least one of the climb command model 1008 or the forward speed command model 1010. For example, based on receiving an enable signal from the left inceptor switch 1002f, the automatic transition function 1003 may automatically determine at least one of the climb signal or forward speed signal to transmit to at least one of the climb command model 1008 or the forward speed command model 1010.

[0088] The turn rate command model 1004 may be configured to output a desired position and / or turn rate command, and may also be configured to calculate a desired heading of the aircraft assumed when the interceptor is returned to the central position (e.g., the detent). The lateral speed command model 1006 may be configured to output a desired position and / or lateral speed command. The climb command model 1008 may be configured to output at least one of a desired altitude, vertical speed, or vertical acceleration command. The forward speed command model 1010 may be configured to output at least one of a desired position, longitudinal speed, or longitudinal acceleration command. In some embodiments, one or more of the command models may be configured to output accelerations generated in response to changes in speed commands. For example, the climb command model 1008 may be configured to output vertical accelerations generated in response to changes in vertical speed commands.

[0089] At least one of the feedforwards 1014 and 1020 (e.g., each) may receive as input one or more desired changes from the corresponding command models 1004, 1006, 1008, 1010 (e.g., desired position, velocity, and / or acceleration), as well as data received from one or more aircraft sensors (e.g., airspeed, vehicle orientation, vehicle load factor, measured acceleration, vehicle mass and inertia, air density, altitude, aircraft mode, etc.), and may be configured to output a corresponding force for at least one (e.g., each) desired change to achieve the desired change. In some embodiments, the feedforwards 1014, 1020 may be configured to determine the corresponding force using a simplified model of aircraft dynamics. For example, based on a known (e.g., stored) or determined mass of the aircraft, the feedforwards 1014, 1020 may be configured to determine a force that causes the aircraft to comply with a desired acceleration command. In some embodiments, the feedforwards 1014, 1020 may be configured to use a model that predicts the amount of drag acting on the vehicle as a function of speed in order to determine the force required to comply with a desired speed command signal.

[0090] At least one of the feedbacks 1012, 1016, 1018, and 1022 (for example, each) may receive as input one or more desired changes from command models 1004, 1006, 1008, and 1010 (e.g., desired position, velocity, and / or acceleration), as well as data received from vehicle detection 1031 indicating vehicle dynamics 1030. For example, the detected vehicle dynamics 1030 may include a representation of the aircraft's physical and / or intrinsic dynamics, and the sensor measurements of vehicle dynamics detection 1031 may capture how the aircraft moves in response to pilot input, propulsion system output, or ambient conditions. Additionally or alternatively, data received from vehicle detection 1031 may include error signals generated by one or more processors based on extrinsic disturbances (e.g., velocity disturbances due to gusts). In some embodiments, feedbacks 1012, 1016, 1018, and 1022 may be configured to generate a feedback force (e.g., in an actuator) based on an received error signal. For example, feedbacks 1012, 1016, 1018, and 1022 may generate a feedback force intended to counteract the effect(s) of an external disturbance. Additionally or alternatively, feedbacks 1012, 1016, 1018, and 1022 may be configured to generate a feedback force based on a modeling error. For example, if an incorrect aircraft mass is input to either feedforward 1014 or 1020, the aircraft may accelerate faster or slower than the desired change. Based on determining the difference between the desired acceleration and the measured acceleration, one or more processors (e.g., included in vehicle detection 1031) may generate an error signal, which may be looped back to feedbacks 1012, 1016, 1018, or 1022 to determine the additional force needed to correct the error.

[0091] In some embodiments, feedbacks 1012, 1016, 1018, and 1022 can be disabled. For example, system 1000 may be configured to operate without feedbacks 1012, 1016, 1018, and 1022 until GPS communication is reconnected, in response to the loss of position and / or ground velocity feedback due to a disruption of Global Positioning System (GPS) communication.

[0092] In some embodiments, feedbacks 1012, 1016, 1018, and 1022 may receive as input a plurality of measurements and a confidence value for at least one (e.g., each) measurement indicating whether the measurement is valid. For example, one or more processors in system 1000 may assign a Boolean value (true / false) to at least one (e.g., each) measurement used in system 1000 to indicate whether the measurement is reliable (e.g., yes) or invalid (e.g., no). Based on one or more processors identifying a measurement as invalid, feedbacks 1012, 1016, 1018, and 1022 may exclude that measurement for further processing. For example, in response to one or more processors identifying a heading measurement as invalid, feedbacks 1012, 1016, 1018, and 1022 may omit subsequent heading measurements when determining the feedback force(s).

[0093] In some embodiments, feedbacks 1012, 1016, 1018, and 1022 may determine one or more feedback forces based on actuator state information received from one or more sensors (e.g., included in vehicle detection 1031). For example, in response to actuator state information indicating an actuator failure, one or more processors of system 1000 may update one or more processes of system 1000 and determine alternative commands to achieve a desired change. For example, one or more processors of system 1000 may adjust one or more models, functions, algorithms, tables, inputs, parameters, thresholds, and / or constraints based on (e.g., in other examples, other aircraft components such as engines or batteries) a change in status (e.g., a failure) of an actuator (or other aircraft component such as an engine or battery). Alternative commands (e.g., yaw, pitch, roll, thrust, or torque) may be determined based on the adjustment(s). Additionally or alternatively, in response to actuator state information indicating that one or more actuators are at their maximum value, one or more processors of system 1000 may update one or more processes of system 1000 (for example, as described above) and determine alternative commands to achieve the desired change.

[0094] The total desired force can be calculated based on the outputs of feedback 1012, 1016, 1018, 1022 and feedforward 1014, 1020. For example, one or more processors of system 1000 may calculate the desired rotation force by summing the outputs of feedback 1012 and feedforward 1014. Additionally or alternatively, one or more processors of system 1000 may calculate the desired lateral force by summing the outputs of feedback 1016 and feedforward 1014. Additionally or alternatively, one or more processors of system 1000 may calculate the desired vertical force by summing the outputs of feedback 1018 and feedforward 1020. Additionally or alternatively, one or more processors of system 1000 may calculate the desired longitudinal force by summing the outputs of feedback 1022 and feedforward 1020.

[0095] At least one of the lateral / directional outer loop allocation 1024 and the longitudinal outer loop allocation 1026 (e.g., each) may be configured to receive as input one or more desired forces and data received from the vehicle detection 1031 (e.g., airspeed, vehicle orientation, vehicle load factor, measured acceleration, vehicle mass and inertia, indicators of working / failed actuators, air density, altitude, aircraft mode, whether the aircraft is in flight or on the ground, etc.). Based on the inputs, the outer loop allocations 1024 and 1026 may be configured to output a roll command, a yaw command, a pitch command, a thrust request, or a combination of different commands / requests in order to achieve one or more desired forces.

[0096] The lateral / directional outer loop allocation 1024 may receive a desired turn rate force and / or a desired lateral force as inputs and may command roll or yaw. In some embodiments, the lateral / directional outer loop allocation 1024 may determine an output based on a determined flight mode. The flight mode may be determined using pilot inputs (e.g., a selected mode for an interceptor) and / or detected aircraft information (e.g., airspeed). For example, the lateral / directional outer loop allocation 1024 may determine the aircraft's flight mode using at least one of determined (e.g., detected or measured) airspeed or inputs received at a pilot interceptor button (e.g., an input that instructs the aircraft to fly according to a particular flight mode). In some embodiments, the lateral / directional outer loop allocation 1024 may be configured to prioritize pilot interceptor button inputs over measured airspeed when determining the flight mode (e.g., the pilot interceptor button is associated with a stronger weight or higher priority than measured airspeed). In some embodiments, the lateral / directional outer loop allocation 1024 may be configured to determine the aircraft's flight mode by blending the determined airspeed and pilot interceptor button input (e.g., using a weighted sum). In hovering flight mode, the lateral / directional outer loop allocation 1024 may achieve a desired lateral force with a roll command (e.g., roll angle, roll rate) and a desired turning rate force with a yaw command. In some embodiments, such as in hovering flight mode, the aircraft may be configured not to accelerate outside a predetermined hovering envelope (e.g., hovering speed range). In forward flight mode (e.g., level flight), the lateral / directional outer loop allocation 1024 may achieve a desired lateral force with a yaw command and a desired turning rate force with a roll command. In forward flight mode, the lateral / directional outer loop allocation 1024 may be configured to determine the output based on the detected airspeed.In the transition between hovering flight mode and forward flight mode, the lateral / directional outer loop allocation 1024 can achieve the desired force using a combination of roll and yaw commands.

[0097] The longitudinal outer loop allocation 1026 may receive a desired vertical force and / or a desired longitudinal force as input and may output at least one of a pitch command (e.g., pitch angle) or a thrust vector request. The thrust vector request may include longitudinal thrust (e.g., a mixture of nacelle tilt and forward propeller thrust) and vertical thrust (e.g., a combined forward and backward thrust). In some embodiments, the longitudinal outer loop allocation 1026 may determine its output based on a determined flight mode. For example, in hovering flight mode, the longitudinal outer loop allocation 1026 may achieve a desired longitudinal force by lowering the pitch attitude and using longitudinal thrust, or by achieving a desired vertical force with vertical thrust. In forward flight mode, the longitudinal outer loop allocation 1026 may achieve a desired longitudinal force with longitudinal thrust (e.g., forward propeller thrust). In cruising flight mode, the longitudinal outer loop allocation 1026 can achieve the desired vertical force by commanding pitch (e.g., increasing pitch attitude) and requesting thrust (e.g., increasing longitudinal thrust).

[0098] The inner loop control law 1028 may be configured to determine a moment command based on at least one of a roll command, yaw command, or pitch command from a lateral / directional outer loop allocation 1024 or a longitudinal outer loop allocation 1026. In some embodiments, the inner loop control law 1028 may depend on detected vehicle dynamics (e.g., from vehicle detection 1031). For example, the inner loop control law 1028 may be configured to compensate for disturbances at attitude and rate levels to stabilize the aircraft. Additionally or alternatively, the inner loop control law 1028 may take into account the period of an eigenmode (e.g., a phugoid mode) affecting the pitch axis and appropriately control the aircraft to compensate for such an eigenmode of the vehicle. In some embodiments, the inner loop control law 1028 may depend on the inertia of the vehicle.

[0099] The inner-loop control law 1028 may determine moment commands using one or more stored dynamic models that reflect the motion characteristics of the aircraft (e.g., the aircraft's aerodynamic damping and / or inertia). In some embodiments, the inner-loop control law 1028 may use a dynamic model (e.g., a low-order equivalent system model) to capture the motion characteristics of the aircraft and determine one or more moments that cause the aircraft to achieve commanded roll, yaw, and / or pitch. Some embodiments may include determining moment commands (e.g., by the inner-loop control law 1028 or other components) based on at least one received command (e.g., a roll command, a yaw command, and / or a pitch command) and a determined (e.g., measured) aircraft state. For example, moment commands may be determined using the difference between a commanded aircraft state and a measured aircraft state. As a further example, moment commands may be determined using the difference between a commanded roll angle and a measured roll angle. As described below, the control allocation 1029 can control the aircraft (e.g., through the flight elements) based on the determined moment command(s). For example, the control allocation 1029 can control (e.g., transmit one or more commands to) one or more electric propulsion systems of the aircraft (e.g., electric propulsion system 602 shown in Figure 6), including tilt actuators(s), electric engines(s), and / or propellers(s). The control allocation 1029 can further control one or more control surfaces of the aircraft (e.g., control surfaces such as flaperons(s), rudder beta(s), ailerons(s), spoilers(s), rudders(s), and / or elevator(s)). Vehicle dynamics 1030 represent the control of different flight elements (e.g., electric propulsion systems and / or control surfaces) and the corresponding effects on the flight elements and aircraft dynamics.

[0100] The embodiment shown in Figure 10 includes both the inner-loop control law 1028 and the outer-loop allocations 1024, 1026, although in some embodiments the flight control system may not include the outer-loop allocations 1024, 1026. Thus, pilot inceptor inputs can generate roll commands, yaw commands, pitch commands, and / or thrust commands. For example, the right inceptor may control roll and pitch, and the left inceptor and / or pedal(s) may control yaw and thrust.

[0101] The control allocation 1029 may accept one or more of the following as inputs: force or moment commands, data received from one or more aircraft sensors, envelope protection limits, scheduling parameters, and optimizer parameters. Based on the inputs, the control allocation 1029 may be configured to determine actuator commands (e.g., thrust(s), torque(s), and / or propeller speed for electric propulsion units) by minimizing an objective function (e.g., solving the optimization problem, such as continuously) which includes one or more primary objectives, such as satisfying commanded aircraft forces and moments (e.g., responding to them, satisfying them, dealing with them, and providing outputs based on them), and one or more secondary objectives, which may include minimizing acoustic noise and / or optimizing battery pack usage.

[0102] In some embodiments, the control allocation 1029 may be configured to calculate limits for individual actuator commands based on the state of the actuator and envelope protection limits. Envelope protection limits may include one or more boundaries in which the aircraft should operate to ensure safe and stable flight. In some embodiments, envelope protection limits may be defined by one or more of the following: speed, altitude, angle of attack, or load factors. For example, envelope protection limits may include one or more bending moments and / or one or more load constraints. In some embodiments, the control allocation 1029 may use the envelope protection limits to automatically adjust one or more control surfaces or control settings. Doing so may protect the aircraft from undesirable scenarios such as stalls or structural strains or failures. Under normal operation, the minimum command limit for a given actuator may include the maximum value of the minimum hardware-based limit and the minimum flight envelope limit, and the maximum command limit may include the minimum value of the maximum hardware-based limit and the maximum flight envelope limit. If an actuator fails, the command limit for that actuator will correspond to its failure mode.

[0103] In some embodiments, the aircraft may include multiple battery packs electrically isolated from each other to power different parts of multiple electric propulsion units. In some embodiments, the control allocation 1029 may include a battery pack energy optimization function. In some embodiments, energy optimization can be achieved by minimizing the use of electric propulsion units connected to battery packs having a lower charge than other battery packs. According to various embodiments, minimizing the use of electric propulsion units connected to battery packs with a lower charge can be achieved by reducing the preferred state of the electric propulsion units in the control allocation optimization objective function (e.g., the optimization problem) and / or increasing the penalty for deviation from the preferred state.

[0104] Additionally or alternatively, the energy optimization function may be configured to modify one or more reference commands (e.g., an attractor command, an ideal actuator command, an ideal effector command, and an ideal combination of commands that cause the aircraft to perform a particular maneuver in the most optimal (e.g., energy-efficient and safe) way). The reference commands may correspond to ideal positions of actuators to achieve one or more desired commands (e.g., optimal commands to achieve a desired motion). In some embodiments, the control allocation 1029 may be configured to determine effector (or actuator) commands based on one or more modified reference commands. For example, at least one (e.g., each) effector may be associated with a reference command, and the control allocation 1029 may generate effector commands for at least one effector such that the deviation between the modified effector reference command and the generated effector command is as small as possible. In some embodiments, the control allocation 1029 may modify one or more reference commands before solving the optimization problem.

[0105] As used herein, the term “effector” means any component configured to produce an effect or result, such as an actuator, control surface, propulsion unit, battery, propeller, engine, or any other aircraft component.

[0106] As used herein, the term “battery pack” means any combination of electrically connected batteries (i.e., battery cells) and may include multiple batteries arranged in series, parallel, or a combination of series and parallel.

[0107] In some embodiments, the control allocation 1029 may include engine thermal management functions for managing (e.g., optimizing) one or more temperatures associated with one or more engines. In some embodiments, engine thermal management may include minimizing the use of electric propulsion units that have a higher temperature than other electric propulsion units. A detailed explanation of engine thermal management is provided in Figures 20 and 21 below.

[0108] In some embodiments, the control allocation 1029 may perform both energy optimization and engine thermal management. In some embodiments, the control allocation 1029 may prioritize engine thermal management over energy optimization. In some embodiments, the control allocation 1029 may prioritize energy optimization over engine thermal management.

[0109] The control allocation 1029 transmits commands to one or more flight elements to control the aircraft. The flight elements are configured to move according to the controlled commands. Various sensing systems and associated sensors as part of the vehicle motion detection 1031 may detect the movement of the flight elements and / or the aircraft's motion and provide information to the feedback 1012, 1016, 1018, 1022, outer loop allocations 1024, 1026, inner loop control rules 1028, and control allocation 1029, which are incorporated into the flight control.

[0110] As described above, the vehicle detection 1031 may include one or more sensors for detecting vehicle dynamics. For example, the vehicle detection 1031 may capture how the aircraft moves in response to pilot input, propulsion system output, or ambient conditions. Additionally or alternatively, the vehicle detection 1031 may detect errors in the aircraft's response based on extrinsic disturbances (e.g., velocity disturbances due to gusts). Furthermore, the vehicle detection 1031 may include one or more sensors for detecting propeller speed, such as a magnetic sensor (e.g., a Hall effect sensor or induction sensor) or an optical sensor (e.g., a tachometer) for detecting the rotor speed (and therefore the propeller speed) of the aircraft engine. The vehicle detection 1031 may include one or more sensors for detecting the nacelle tilt angle (e.g., the propeller rotation axis angle between the lift configuration (e.g., Figure 2) and the forward thrust configuration (e.g., Figure 1)), such as a magnetic sensor (e.g., a Hall effect sensor or induction sensor). The vehicle detection system 1031 may include one or more sensors configured to detect engine torque and / or thrust, such as one or more current sensors or voltage sensors, strain gauges, load cells, and / or propeller vibration sensors (e.g., accelerometers).

[0111] The vehicle detection system 1031 may include one or more sensors configured to detect vehicle dynamics, such as acceleration sensors and / or pitch orientation sensors (e.g., accelerometers, 3-axis accelerometers, gyroscopes, 3-axis gyroscopes, and / or tilt position sensors for determining engine angle), and airspeed sensors (e.g., Pitot tube sensors). The vehicle detection system 1031 may further include one or more inertial measuring units (IMUs) for determining aircraft conditions based on these measurements. Aircraft conditions may refer to forces experienced by the aircraft, orientation of the aircraft, position of the aircraft (e.g., altitude), and / or motion of the aircraft. For example, the aircraft state may include at least one of the following: the aircraft's position (e.g., yaw angle, roll angle, pitch angle, and / or any other orientation across one or two axes), the aircraft's speed, the aircraft's angular rates (e.g., roll rate, pitch rate, and / or yaw rate), and / or the aircraft's acceleration (e.g., longitudinal, lateral, and / or vertical acceleration), or any physical properties of the aircraft or one of its components. In some embodiments, the vehicle detection 1031 may include an inertial navigation system (INS) and / or atmospheric data and / or attitude heading reference system (ADAHRS). The inertial navigation system (INS) and / or atmospheric data and attitude heading reference system (ADAHRS) may include one or more inertial measurement units (IMUs) and corresponding sensors (e.g., accelerometers, gyroscopes, 3-axis gyroscopes, and / or 3-axis accelerometers). In some embodiments, the INS and / or ADAHRS may filter and / or process sensor measurements to determine aircraft conditions (e.g., acceleration or angular rate). For example, in some embodiments, the INS and / or ADAHRS may determine the angular rate based on gyroscope measurements and the acceleration based on accelerometer measurements.

[0112] DCPS1033 may receive inputs from control allocation 1029 and vehicle detection 1031. For example, DCPS1033 may receive at least one torque command, one or more limits (e.g., envelope protection limit, engine torque rate limit, HV channel / HV bus protection torque limit, etc.), or measured vehicle dynamics (e.g., measured revolutions per minute (RPM), measured voltage, etc.). DCPS1033 may be configured to dynamically generate one or more modified torque commands (e.g., dynamically varied torque commands) by modifying one or more torque commands based on the inputs. System 1000 may be configured to transmit one or more dynamically modified torque commands to one or more engines of the aircraft. The disclosed embodiments may improve engine response and enhance aircraft stability and safety.

[0113] Figure 11 illustrates a block diagram of an exemplary flight control system 1100, including an energy optimization function, according to various embodiments. In particular, considering that these functionalities are implemented frequently (e.g., constantly, continuously), it is understood that the steps of the exemplary methods depicted in Figure 11 would be impossible, or at least impractical, to effectively implement by a human user while the aircraft is in flight (including takeoff or landing) and / or dynamically based on (e.g., in response to) received signals (e.g., aircraft sensors, pilot input devices). Generally, it can be understood that any / all steps of the exemplary methods in Figure 11 may be performed or executed by at least one processor (e.g., FCS, system 1000), such as according to one or more instructions stored in a computer-readable medium (e.g., non-temporary computer-readable medium). In some embodiments, the flight control system may include an energy optimization function configured to update one or more initial reference commands (e.g., reference torque, reference torque command, reference speed, reference speed command, reference power, reference power command, reference current, and / or reference current command). In some embodiments, the flight control system may be configured to recalculate forces and / or moments based on one or more updated initial reference commands, calculate allocations (as illustrated, for example, with respect to control allocations in Figure 10), satisfy (e.g., respond to, satisfy, deal with, and based on) the recalculated forces and / or moments, and minimize the difference from one or more updated initial reference commands. Figure 11 illustrates an example of energy optimization associated with torque, but it should be further understood that Figure 11 may, alternatively or additionally, be associated with any parameters related to aircraft effectors such as speed, power, and / or current.

[0114] In 1110, the energy optimization function is based on one or more initial reference commands 1102 (e.g., reference torque commands, τ ref0The energy optimization function may update one or more initial reference directives 1102 based on battery information 1104 (e.g., available energy (e.g., available energy) from the BMS), remaining discharge time (e.g., of one or more battery packs, one or more crosslinks, one or more independent high-voltage buses, one or more high-voltage channels), energy state (SOE), charge state (SOC), power state (SOP), health state (SOH), fault conditions (e.g., short circuit or overcurrent, whether the battery is active / functioning), or temperature state (SOT) of at least one (e.g., each) battery pack. For example, if one or more first rotors (e.g., engines, motors, EPUs) are connected to a first battery pack having lower available energy (e.g., compared to the average available energy of all available (e.g., active, functional) battery packs compared to other battery packs), the energy optimization function may update one or more first initial reference commands of one or more initial reference commands 1102 associated with one or more first rotors to be within a first range before performing the allocation function (e.g., solving the allocation problem as described with respect to the control allocation 1029 in Figure 10). Additionally or alternatively, if one or more second rotors are connected to a second battery pack having higher available energy (e.g., compared to the average available energy of all available (e.g., active, functional) battery packs compared to other battery packs), the energy optimization function may update one or more second initial reference commands of one or more initial reference commands 1102 associated with one or more second rotors to be within a second range higher than the first range before performing the allocation function. In some embodiments, the energy optimization function may update one or more initial reference commands 1102 based on battery information 1104 received from one or more battery management systems associated with a plurality of battery packs, based on the difference in state between at least two of the plurality of battery packs (e.g., between each battery pack).In some embodiments, the energy optimization function may be configured to calculate one or more updated initial reference torque commands, as follows: TIFF2026525201000002.tif176158

[0115] As discussed herein, available energy (e.g., available energy) may be determined by at least one BMS or flight control system based on battery information (e.g., measured information), e.g., remaining discharge time (e.g., of one or more battery packs, one or more crosslinks, one or more independent high-voltage buses, one or more high-voltage channels), energy state (SOE), state of charge (SOC), power state (SOP), health state (SOH), fault conditions (e.g., short circuit or overcurrent, whether the battery is active / functioning), or temperature state (SOT) of at least one (e.g., each) battery pack.

[0116] In some embodiments, the energy optimization function may determine one or more initial reference commands 1102. Additionally or alternatively, the energy optimization function may receive one or more initial reference commands 1102 from a flight control law (as considered, for example, with respect to system 1000). For example, a flight control system may use a flight control law to determine one or more initial reference commands 1102. In some embodiments, the flight control system may include a lookup model configured to determine one or more reference commands (multiple) (e.g., initial reference torque command, initial reference power command, initial reference velocity command, initial reference current command), reference effector position, reference force, reference moment, and / or one or more Jacobians (e.g., Jacobian matrices) based on one or more lookup tables. For example, a lookup model may refer to one or more lookup tables containing scheduling parameters, and may use one or more current aircraft conditions (e.g., vehicle dynamics (e.g., Figure 10), ambient conditions, flight conditions (e.g., altitude, airspeed, acceleration, angle of attack, torque of one or more propellers, inclination angle and / or orientation of one or more propellers, whether drive lift is active, acceleration, load coefficient drive flaperon schedule, position / motion of control surfaces, flight phase, speed of one or more propellers, and / or engine torque of one or more engines) and the status and / or health of at least one (e.g., each) aircraft component (e.g., temperature, time, energy, and / or availability of battery packs, propulsion units, engines, high-voltage buses, and / or channels) to determine the baseline and Jacobean values.Scheduling parameters may include one or more of the following schedulings: EPU scheduling (e.g., adjusting rotor / propeller speed and direction in different phases of flight), battery management (e.g., changing power consumption from different battery packs to ensure optimal performance and lifespan), control surface motion (e.g., adjusting the motion of flaperons, elevators, rudders, or any other control surfaces in different phases of flight), power distribution (e.g., ensuring critical systems receive priority power during any high-demand situations), timing of inputs to manage the aircraft's attitude, altitude, and / or trajectory, thermal management (e.g., scheduling the activation of cooling systems to manage the temperature of critical components), data processing (e.g., timing of data collection), emergency response protocols (e.g., scheduling the activation of emergency procedures in response to certain triggers or conditions), autonomous system scheduling, or any scheduling to describe how different components of the aircraft are managed or adjusted for various aircraft conditions to ensure optimal performance, safety, and efficiency.

[0117] In some embodiments, the energy optimization function may be configured to clip one or more initial reference commands. For example, the energy optimization function may automatically clip one or more initial reference commands based on at least one of one or more envelope protection limits or status information (e.g., active, inactive, healthy, energy level, available) associated with one or more aircraft components (e.g., engine, battery, bus, channel). In some embodiments, the energy optimization function may clip the initial rotor reference torque so that it is within the envelope protection range (e.g., below the maximum envelope protection limit, above the minimum envelope protection limit). In some embodiments, the energy optimization function may receive one or more clipped initial reference commands (e.g., clipped by the flight control system using flight control laws).

[0118] In 1120, the flight control system may recalculate (e.g., adjust) one or more force or moment (FM) commands. For example, the flight control system may, based on one or more initial force or moment commands 1122 (e.g., inputs to control allocation 1029 in Figure 10, Fcmd0) and one or more outputs of 1110 (e.g., one or more changes (e.g., delta) in the initial reference command(s) 1106, the difference between the updated reference command(s) and the initial reference command(s), Δτ ref Based on this, one or more force or moment commands may be recalculated. In some embodiments, the flight control system may recalculate one or more force or moment commands as follows: TIFF2026525201000003.tif137158

[0119] In step 1130, the flight control system may perform a control allocation (e.g., resolving the allocation, a step performed by the control allocation 1029 in Figure 10) based on the allocation input 1132, the updated initial reference command(s) 1108, and the recalculated force / moment command 1124. In some embodiments, as a result of performing step 1130 (e.g., based thereon), the flight control system may perform one or more effector commands 1134 (e.g., the output of the control allocation 1029 in Figure 10, τ cmd It may output (etc.). For example, based on the received allocation input 1132 (e.g., the output of the inner loop control law 1028, the output of the outer loop allocation 1026, and / or the input to the control allocation 1029 in Figure 10, data received from one or more aircraft sensors, envelope protection limits, scheduling parameters, and / or optimizer parameters), the flight control system (e.g., the control allocation 1029 in Figure 10) may resolve the allocation such that the recalculated force and / or moment command 1124 is satisfied (e.g., completely within the threshold margin) and the difference between the updated initial reference command(s) 1108 and one or more output effector commands 1134 is minimized.

[0120] In some embodiments, the energy optimization function may be considered a secondary objective to at least one other objective or function (e.g., implemented by the flight control computer and / or system 1000). For example, energy optimization may have a lower priority than engine thermal management, which may be implemented by the engine thermal management function, as considered below. In some embodiments, the time scale of the electric engine may be shorter than the time scale of battery energy optimization. For example, an FCC or other device implementing different functions disclosed herein may optimize electric engine temperature parameters over a time period shorter than the time period for optimizing battery energy parameters. In some embodiments, when the energy optimization function and the engine thermal management function apply adjustments to a reference value in the same direction, the larger of the two adjustments may be applied. In some embodiments, when the energy optimization function and the engine thermal management function produce opposing outputs, the flight control system may associate the reference value associated with the engine thermal management function with a greater weight than the reference value associated with the energy optimization function (e.g., only the reference value associated with the thermal management function may be considered). In some embodiments, the flight control system may be configured to dynamically prioritize (e.g., balance) energy optimization or engine thermal management based on the current phase of flight, the temperature of one or more aircraft components (e.g., engines, batteries), and / or remaining energy.

[0121] In some embodiments, the flight control system may include a battery thermal management function, which may be configured similarly to an engine thermal management function. For example, the battery thermal management function may be configured to normalize the temperature across different battery components, which may have different time constants and temperature limits, using a time-based metric for temperature.

[0122] Figure 12 is a flowchart illustrating exemplary method 1200 for energy optimization of a particular machine (e.g., an aircraft) according to some embodiments of the present disclosure, thereby improving aircraft technology (e.g., aircraft safety, energy efficiency, range capability, payload capacity, structural integrity). The steps of method 1200 may be carried out, for example, by the system 1000 of Figure 10 being carried out on or using in other ways any aircraft component associated with any of Figures 1 to 8, 9A to 9E, 10, 11, 13A to 13D, or 14 to 25, or any flight control computer (e.g., a method performed by a computer) or flight control system. For example, an aircraft flight control computer may be configured to carry out one or more steps of method 1200.

[0123] The steps of Method 1200 may be activated or modified as needed, based on the aircraft's flight conditions, maneuvering, or operating requirements. It is understood that the illustrated Method 1200 may be modified by changing the order of the steps and by including additional steps. It is also understood that the complexity of Method 1200 would be impossible, or at least extremely impractical, to implement effectively by a human user, especially considering that these functionalities are implemented (e.g., in real time) while the energy and battery conditions used by the aircraft are consistently changing, particularly while the aircraft is in flight (including takeoff or landing). Furthermore, energy optimization (e.g., using Method 1200) improves the aircraft's safety and range capability (e.g., how much energy is available to power the aircraft) as well as the lifespan of different components (e.g., batteries) without requiring the direct involvement of a user (e.g., a pilot). In some embodiments, the aircraft of System 1000 may include an electric aircraft including multiple effectors, including multiple electric propulsion units (EPUs) and multiple battery packs powering the multiple electric propulsion units.

[0124] In step 1202, system 1000 may determine one or more desired commands for the electric aircraft. For example, one or more desired commands may be one or more force or moment commands (e.g., inputs to control allocation 1029 in Figure 10), a desired position and / or turn rate command (e.g., output of turn rate command model 1004 in Figure 10), a desired position and / or lateral velocity command (e.g., output of lateral velocity command model 1006 in Figure 10), one or more desired altitude, vertical velocity, or vertical acceleration commands (e.g., output of climb command model 1008 in Figure 10), one or more desired position, longitudinal velocity, or longitudinal acceleration commands (e.g., output of forward velocity command model 1010 in Figure 10), alternative commands to achieve a desired change (e.g., based on feedback as considered in Figure 10), one or more desired total force commands (e.g., feed in Figure 10) The commands may include at least one of the following commands associated with a desired movement of the aircraft (e.g., desired by the aircraft's pilot, desired by the flight control system, desired by the autopilot system): one or more moment commands (e.g., determined by the inner-loop control law 1028 in Figure 10), linear commands, angular acceleration commands, pitch commands, roll commands, yaw commands, or one or more signals (e.g., pilot inputs from pilot input devices, inputs from the autopilot system, signals generated by the flight control system or computer).

[0125] In step 1204, the system 1000 may determine at least one reference command (e.g., an initial reference command) based on one or more desired commands and one or more aircraft conditions. The reference command may include attractors (e.g., ideal flight conditions, desired state, ideal actuator / effector position / velocity / torque / current / power) to achieve one or more desired commands. For example, the system 1000 may be configured to constantly monitor the aircraft conditions or state (e.g., using a feedback control system) and may generate or adjust control commands to correct any deviation from the reference command (e.g., to adjust force / moment commands, to adjust aircraft components such as actuators, effectors, propulsion units, control surfaces, etc.). In some embodiments, the system 1000 may determine at least one reference command based on one or more aircraft conditions. The aircraft conditions may include one or more of the following: (e.g., vehicle dynamics, ambient conditions, flight conditions (e.g., altitude, airspeed, angle of attack, torque of one or more propellers, inclination angle of one or more propellers, whether power lift is active, acceleration, load coefficient drive flaperon schedule, flight phase, speed of one or more propellers, engine torque of one or more engines), and / or the status and / or health of at least one (e.g., each) aircraft component (e.g., battery, propulsion unit, high-voltage bus, and / or channel availability). In some embodiments, determining at least one reference command may include determining at least one initial (e.g., predetermined, stored) reference command based on the current aircraft conditions using a plurality of predetermined lookup tables. In some embodiments, the system 1000 may determine at least one reference command for at least one (e.g., each) actuator, effector, or both. In some embodiments, the system 1000 may determine at least one reference command for at least one (e.g., each) force, moment, or both. In some embodiments, the at least one reference command may include at least one of a position command, trim command, torque command, power command, current command, or speed command.

[0126] In step 1206, system 1000 may monitor the energy state of multiple battery packs. The energy state may be represented by and / or include one or more battery parameters (e.g., related to energy consumed, remaining energy, available energy, usable energy, and / or energy to be consumed) that indicate the measured, predicted, and / or capability of the battery packs. In some embodiments, system 1000 may receive at least one (e.g., each) battery pack associated battery parameter from one or more BMSs associated with the multiple battery packs. The battery parameter may include one or more of the following: available energy, remaining discharge time (e.g., of one or more batteries, one or more crosslinks, one or more independent high-voltage buses, one or more high-voltage channels), energy state (SOE), charge state (SOC), power state (SOP), health state (SOH), fault condition (e.g., short circuit or overcurrent, whether the battery is active / functional / malfunctioning / failed), or at least one (e.g., each) battery pack temperature state (SOT). In some embodiments, system 1000 may calculate at least one of available energy or remaining discharge time based on other battery information. In some embodiments, the battery information may include the difference in power consumption (e.g., of EPUs) between multiple EPU engines. In some embodiments, system 1000 may be communicatively coupled to one or more battery management systems ("BMS") of an aircraft (e.g., via physical connections such as bus and / or channel systems, or via digital communication interfaces). In some embodiments, an aircraft may include a single BMS configured to manage all battery packs on the aircraft. In some embodiments, at least one (e.g., each) battery pack may include its own BMS.

[0127] In some embodiments, system 1000 may also monitor the status of at least one (e.g., each) propulsion unit among a plurality of electric propulsion units. For example, system 1000 may receive EPU information associated with the EPU (e.g., EPU status, engine status, and / or propeller status) from at least one (e.g., each) EPU among a plurality of EPUs (e.g., via a digital communication interface, via the EPU inverter). EPU information may include one or more of the following: the temperature of at least one component (e.g., engine, propeller) of at least one EPU; one or more times associated with the temperature of at least one component (e.g., the remaining time of the engine at the current torque rating, such as continuous torque limit (e.g., 2204 in Figure 22), engine takeoff torque limit (e.g., 2205 in Figure 22), engine emergency torque limit (e.g., 2206 in Figure 22), engine overheat torque limit (e.g., 2207 in Figure 22), predicted duration at maximum takeoff torque (e.g., 2208 in Figure 22), or predicted duration at maximum emergency torque (e.g., 2209 in Figure 22); or fault conditions (e.g., whether the EPU and / or components of the EPU are active and responding to commands). In some embodiments, monitoring the energy state of multiple battery packs may include predicting the energy information (e.g., capacity, capability) of the battery packs for one or more future times (e.g., multiple points in time) as will be further considered below. In some embodiments, the system 1000 may use the predicted energy information to influence aircraft operation to extend the aircraft's flight range and / or cause multiple battery packs to be discharged simultaneously or nearly simultaneously (e.g., within acceptable limits) (or nearly discharged, such as being discharged below a certain threshold).

[0128] In step 1208, the system 1000 may adjust at least one reference command (e.g., an initial reference command) based on the monitored energy states of multiple battery packs. In some embodiments, adjusting at least one reference command may involve clipping at least one reference command using one or more command limits (e.g., envelope protection limits, minimum and / or maximum limits associated with effectors / actuators, and at least one command parameter (e.g., position, velocity, torque, power, current)) such that the clipped reference command is within the command limits (e.g., down, up).

[0129] In some embodiments, adjusting at least one reference command may involve calculating one or more estimated reference commands for at least one (e.g., each) high-voltage bus / channel among a plurality of high-voltage buses / channels. In some embodiments, the one or more estimated reference commands may include at least one of one or more power reference commands, one or more position reference commands, one or more trim reference commands, one or more torque reference commands, one or more speed reference commands, or one or more current reference commands. In some embodiments, calculating one or more estimated reference commands may be based on one or more aircraft component connections (e.g., high-voltage architecture connections between one or more engines and one or more high-voltage channels and / or buses, connections between battery packs, connections between one or more engines and one or more battery packs). For example, system 1000 may use a connection matrix containing connection information associated with the aircraft's high-voltage architecture to calculate one or more estimated reference commands for at least one (e.g., each) high-voltage bus / channel. In some embodiments, the connection matrix may map at least one (e.g., each) engine to an HV bus / channel. In some embodiments, the connection matrix may describe connections between aircraft components (e.g., between engine / rotor and HV channel / bus, between battery pack and engine / rotor, between battery packs). In some embodiments, the connection matrix may be stored in the flight control law. In some embodiments, the system 1000 may dynamically update the connection matrix based on dynamic updates to connections (e.g., based on failed connections, based on identifying failures in one or more components connected to the high-voltage architecture).

[0130] In some embodiments, adjusting at least one reference command may involve determining one or more ratios (e.g., energy ratio, power ratio, torque ratio, speed ratio, current ratio) of at least one (e.g., each) high-voltage channel or high-voltage bus. For example, the ratio may include a normalized ratio representing the difference between estimated reference commands (e.g., total power command, total position command, total trim command, total torque command, total speed command, and / or total current command) between different buses / channels. In some embodiments, system 1000 may determine the ratio for each high-voltage bus / channel based on the average of the total commands (e.g., total power command, total current command, total position command, total trim command, total torque command, and / or total speed command) of the high-voltage buses / channels. For example, system 1000 may determine the power ratio for each high-voltage bus as follows: TIFF2026525201000004.tif150155

[0131] In some embodiments, adjusting at least one reference command may include determining the energy ratio of at least one (e.g., each) high-voltage bus / channel. For example, determining the energy ratio may include determining the total available energy associated with at least one (e.g., each) high-voltage bus / channel. Additionally or alternatively, system 1000 may normalize the total available energy based on the average of the energy available on at least one (e.g., each) bus / channel. In some embodiments, system 1000 may determine the energy ratio per high-voltage bus as follows: TIFF2026525201000005.tif105129

[0132] In some embodiments, adjusting at least one reference command may involve updating one or more reference commands to optimize the remaining discharge time across all high-voltage buses / channels. For example, system 1000 may adjust at least one reference command (e.g., initial reference command, clipped reference command) based on one or more determined ratios (e.g., power ratio, energy ratio, current ratio, torque ratio, speed ratio) and connection matrices. In some embodiments, system 1000 may adjust at least one reference command based on a relationship between at least two ratios. For example, system 1000 may adjust at least one reference command based on a relationship (e.g., ratio) between one or more determined power ratios and one or more determined energy ratios. In some embodiments, system 1000 may determine (e.g., optimize) the remaining discharge time for at least one (e.g., each) high-voltage bus / channel based on one or more of the at least one reference command, one or more determined ratios, or connection matrices. For example, system 1000 may update one or more reference commands to balance the remaining discharge time across all buses / channels (e.g., equalize within a threshold), as follows: TIFF2026525201000006.tif153136

[0133] In some embodiments, determining (e.g., optimizing) the remaining discharge time may include adjusting at least one reference command such that the remaining discharge time for at least one (e.g., each) high-voltage bus / channel is the same. Additionally or alternatively, determining (e.g., optimizing) the remaining discharge time may include adjusting at least one reference command such that the remaining discharge time for at least one (e.g., each) high-voltage bus / channel is within a predetermined range.

[0134] In some embodiments, the system 1000 may adjust a first reference command associated with the first engine and a second reference command associated with the second engine when the monitored energy state indicates that the first engine has lower power consumption than the second engine. For example, the system 1000 may increase the torque or power command associated with the first reference command in relation to the torque or power command associated with the second reference command.

[0135] In some embodiments, the system 1000 may adjust at least one reference command based on the monitored state of at least one (e.g., each) propulsion unit among a plurality of electric propulsion units. For example, based on received EPU information, the system 1000 may adjust (e.g., generate, update) one or more control commands (e.g., power command, position command, trim command, torque command, speed command, current command) transmitted to one or more EPUs (or their components, such as propellers or engines).

[0136] In some embodiments, system 1000 may adjust at least one reference command based on further predictions of power consumption (e.g., power required or used by one or more components of an aircraft). For example, system 1000 may use simulation data (e.g., stored) or historical data to perform power consumption predictions based on one or more aircraft conditions. In some embodiments, the simulation and / or historical data may show the power consumption of at least one (e.g., each) engine associated with different aircraft conditions. In some embodiments, the simulation or historical data may include one or more of the following: flight test data, weather data, route data, expected discharge data, distance data, or battery health data.

[0137] In some embodiments, system 1000 may clip at least one adjusted reference command. For example, system 1000 may clip at least one reference command using one or more command limits (e.g., envelope protection limits, minimum and / or maximum limits associated with effectors / actuators, and command parameters (e.g., position, velocity, torque, power, current)) such that the clipped adjusted reference command is within command limits (e.g., down, up). In some embodiments, the adjusted at least one reference command in steps 1210 and 1212 may refer to the clipped adjusted at least one reference command.

[0138] In step 1210, the system 1000 may generate control commands for multiple effectors based on at least one adjusted reference command. In some embodiments, generating control commands for multiple effectors may include calculating the difference between forces and moments based on at least one adjusted reference command. In some embodiments, generating control commands may include recalculating forces or moments (for example, as shown in Figure 11) based on at least one adjusted reference command. For example, the system 1000 may recalculate forces and / or moments by determining changes in forces and / or moments, as follows: TIFF2026525201000007.tif133161

[0139] In some embodiments, generating control commands for multiple effectors may include calculating actuator limits based on at least one adjusted reference command. In some embodiments, generating control commands for multiple effectors may include generating multiple control commands and selecting one or more control commands that are closest to at least one reference command.

[0140] In some embodiments, system 1000 may generate control commands for multiple effectors based on one or more temperatures associated with one or more aircraft components (e.g., engines, battery packs). For example, system 1000 may generate control commands for multiple effectors based on one or more temperatures associated with one or more engines so that one or more temperatures remain within one or more predetermined ranges (e.g., continuous range, takeoff / landing range, emergency range). For example, system 1000 may generate control commands to adjust weights or other parameters that reduce the torque commanded for one or more engines. In some embodiments, system 1000 may generate control commands for multiple effectors so that one or more temperatures associated with one or more aircraft components do not exceed predetermined temperatures (e.g., operating limits, immediate fault limits, warning limits).

[0141] In some embodiments, system 1000 may generate control commands for multiple effectors based on priorities associated with energy optimization and temperature control. For example, system 1000 may determine a first remaining time associated with available energy on the aircraft and a second remaining time associated with reaching a temperature limit. Based on the first remaining time being greater than the second remaining time, system 1000 may prioritize temperature control over energy optimization. Additionally or alternatively, based on the second remaining time being greater than the first remaining time, system 1000 may prioritize energy optimization over temperature control.

[0142] In step 1212, system 1000 may control a plurality of effectors according to the generated control commands to satisfy (e.g., respond to, satisfy, deal with, and based on) a desired command of the electric aircraft. For example, system 1000 may transmit at least one (e.g., each) generated control command to its respective effector to move the effector according to the controlled command.

[0143] In some embodiments, Method 1200 can result in optimized energy use so as to maximize the amount of energy available to the aircraft for use. In some embodiments, Method 1200 can result in all battery packs running out of energy simultaneously. In effect, Method 1200 can result in maximizing the available flight range and can also enhance safety by avoiding dangerous situations where one or more batteries run out of energy before the others, making it impossible to power their respective EPUs.

[0144] Figure 13A illustrates an exemplary battery pack failure scenario without energy optimization, according to several embodiments, and Figure 13B illustrates an exemplary battery pack failure scenario with energy optimization (using, for example, system 1100 and / or method 1200). Both figures illustrate the bus energy and power consumption of the aircraft's three buses (each connected to two batteries) at various points in the flight or trip (e.g., start, mid-flight distance, end) in this exemplary scenario, with the battery pack connected to bus 1 failing at the midpoint. As shown in Figure 13A, without energy optimization, the same amount of power is drawn from each bus (e.g., bus 1, bus 2, bus 3) even after the battery failure in bus 1, and bus 1 runs out of energy at approximately 3 / 4 of the distance of the entire trip. Conversely, with energy optimization, as shown in Figure 13B, after the battery pack connected to bus 1 fails, the energy optimization function adjusts the amount of power drawn from each bus to account for the battery pack failure. In this way, the aircraft can reach approximately 11 / 12 of its mission distance and deplete all battery packs simultaneously. By implementing energy optimization, such as by using the techniques described with respect to Figure 11 or Figure 12, the aircraft may be able to extend its flight range by distributing the energy or power drawn from its battery packs.

[0145] Figure 13C illustrates an exemplary battery pack failure scenario without energy optimization in several embodiments, and Figure 13D illustrates an exemplary battery pack failure scenario with energy optimization (e.g., using System 1100 and / or Method 1200). Both figures illustrate the bus energy and power consumption of three buses (each connected to two batteries) of an aircraft at various flight distances, in this exemplary scenario, at the midpoint, a battery pack connected to one of the buses fails. As shown in Figure 13C, without implementing energy optimization and starting with 50% buffer energy in addition to the nominal energy (e.g., 50% of the nominal energy), even after a battery failure in bus 1, the same amount of power is drawn from each bus, with one of the buses depleted and two of the buses still having 50% energy remaining at the end of the mission. Conversely, Figure 13D implements energy optimization, starting with 10% buffer energy in addition to the nominal energy, and the amount of power drawn from each bus is adjusted to account for the battery pack failure. In this way, the aircraft can reach the end of the journey, completely depleting the energy from all three buses, using less buffer energy than would be possible without energy optimization. By implementing energy optimization, such as by using the techniques described with respect to Figure 11 or Figure 12, less buffer energy may be required to complete the mission.

[0146] In some embodiments, the energy optimization function may be configured to receive battery information (e.g., charge state, energy state, power state, health state, temperature state, available energy, etc.) for at least one (e.g., each) battery from the battery management system (BMS) of the flight control system. The charge state (SOC) may refer to the available capacity of the battery pack relative to its rated capacity. For example, the charge state is obtained based on the open-circuit voltage (OCV) of the battery pack, where OCV is the quiescent voltage of the battery pack (e.g., based on a battery pack with no current flow for a set period of time). The energy state (SOE) may be a measure of the remaining discharge energy of the battery. The power state (SOP) may be the maximum power capacity that the battery can deliver over a given period of time. The health state (SOH) may refer to the overall condition of the battery compared to its ideal condition (e.g., degradation level). The temperature state (SOT) may refer to the current temperature of the battery. Based on received battery information relating to at least one (e.g., each) battery, as well as the high-voltage architecture of the flight control system, the energy optimization function may set or adjust reference points (e.g., reference commands) for the flight control system to implement control allocations.

[0147] Figures 14 to 18 illustrate exemplary scenarios for implementing energy optimization functions (e.g., all or part of system 1100 and / or method 1200) according to several embodiments. In each scenario, HV bus 1 is associated with batteries 1 and 4, as well as engines 1, 4, 9, and 12, respectively, connected to HV channels 1 and 4 (not shown). For example, battery 1 is connected to engines 1 and 12 via HV channel 1, and battery 4 is connected to engines 4 and 9 via HV channel 4. HV bus 2 is associated with batteries 2 and 5, as well as engines 2, 5, 8, and 11, respectively, connected to HV channels 2 and 5 (not shown). For example, battery 2 is connected to engines 2 and 11 via HV channel 2, and battery 5 is connected to engines 5 and 8 via HV channel 5. HV bus 3 is associated with batteries 3 and 6, as well as engines 3, 6, 7, and 10, respectively, connected to HV channels 3 and 6 (not shown). For example, battery 3 is connected to engines 3 and 10 via HV channel 3, and battery 6 is connected to engines 6 and 7 via HV channel 6.

[0148] The "battery index" (e.g., battery indices 1402, 1502, 1602, 1702, 1802 in Figures 14-18) may indicate the available energy associated with each battery pack (e.g., batt1, batt2, batt3, batt4, batt5, batt6). The "HV bus energy index" (e.g., HV bus energy indices 1404, 1504, 1604, 1704, 1804 in Figures 14-18) may indicate the available energy associated with each HV bus (e.g., bus1, bus2, bus3) and / or HV channel (e.g., ch1, ch4 in Figure 18). The "engine index" (e.g., engine indices 1406, 1506, 1606, 1706, 1806 in Figures 14-18) may represent the initial reference power command (e.g., "input reference") and the adjusted reference power command (e.g., "output reference") associated with each engine (e.g., EE1-EE12). The "HV bus power index" (e.g., HV bus power indices 1408, 1508, 1608, 1708, 1808 in Figures 14-18) may represent the initial reference command (e.g., "input reference") and the adjusted reference command (e.g., "output reference") for power consumption associated with each HV bus (e.g., bus1, bus2, bus3) and / or HV channel (e.g., ch1, ch4 in Figure 18).

[0149] Figure 14 illustrates nominal scenario 1400, in which each battery pack of batteries 1-6 has equivalent available energy (indicated by, for example, battery index 1402) and all HV channels / buses have equivalent available energy (indicated by HV bus energy index 1404). In scenario 1400, the initial reference command input (e.g., "input reference") to the energy optimization function may be the same as the reference command output / adjusted by the energy optimization function (e.g., "output reference"), as indicated by engine index 1406 and HV bus power index 1408.

[0150] Figure 15 illustrates battery failure scenario 1500, in which battery batt1 fails and all other batteries (e.g., batt2-batt6) have equivalent available energy (indicated by battery index 1502). HV bus bus1 associated with the failed batt1 has approximately half the available energy of HV buses bus2 and bus3 respectively (indicated by HV bus energy index 1504). As indicated by engine index 1506, the energy optimization function adjusts the reference command associated with engines EE1, EE4, EE9, and EE12 (powered only by battery batt2 via HV bus bus1) to a low value (e.g., "output reference" is approximately 0.6) and adjusts the reference command associated with engines EE2, EE3, EE5-EE8, EE10, and EE11 to a high value (e.g., "output reference" is approximately 1.2). As shown in the HV bus power index 1508, the energy optimization function may adjust the reference command associated with HV bus bus1 connected to a failed battery pack batt1 to a lower value (e.g., "output reference" is 2.4) and increase the respective reference commands associated with HV buses bus2 and bus3 to a higher value (e.g., "output reference" for both bus2 and bus3 is 4.8) in order to reduce power consumption from bus bus1 and increase power consumption from buses bus2 and bus3. In this way, HV bus bus1 may be prevented from exhausting its energy before HV buses bus2 and bus3.

[0151] Figure 16 illustrates engine failure scenario 1600 in which engine EE1 has failed. Each battery pack from batt1 to batt6 has equivalent available energy (indicated by battery index 1602), and all HV buses from bus1 to bus3 have equivalent available energy (indicated by HV bus energy index 1604). As indicated by engine index 1606, the output reference power command associated with engine EE1 (e.g., "output reference") collapses to 0, and the energy optimization function adjusts the reference power commands associated with engines EE4, EE9, and EE12 to be higher (e.g., "output reference" associated with 4, 9, and 12), thereby increasing the torque / speed / usage associated with engines EE4, EE9, and EE12. The energy optimization function reduces the torque / speed / usage associated with engines EE2, EE3, EE5-EE8, EE10, and EE11 by adjusting the reference power commands associated with engines EE2, EE3, EE5-EE8, EE10, and EE11 to a lower value (e.g., the "output references" associated with 2, 3, 5-8, 10, and 11). As shown in HV bus power index 1608, the energy optimization function may adjust the reference commands so that the output reference commands associated with each bus are approximately the same (e.g., the "output references" for buses 1-3), even though HV bus bus 1 connected to the failed engine EE1 is used only to supply power to engine EE12. As shown in scenario 1600, in response to an engine failure, the energy optimization function may adjust the reference commands associated with the engines so that the remaining energy or remaining discharge time for each HV bus is approximately the same (e.g., the difference is within a threshold).

[0152] Figure 17 illustrates HV channel failure scenario 1700, in which HV channel 1 of bus b1 fails. Due to the failure of HV channel 1, the available energy of battery batt1 may drop to 0 (indicated by battery index 1702), and HV bus batt1 may only have half the available energy of bus bus2 or bus3, respectively (indicated by HV bus energy index 1704). As indicated by engine index 1706, the failure of HV channel 1 collapses the reference commands associated with engines EE1 and EE12 to 0, which may automatically reduce the reference command associated with bus bus1 (indicated by HV bus power index 1708), so the energy optimization function may not need to adjust the engine reference commands. As shown in scenario 1700, the effect of losing the battery may be mitigated (e.g., balanced) by the effect of losing the two engines.

[0153] Figure 18 illustrates a crosslink failure scenario 1800 in which the crosslink (e.g., HV bus but1) connecting battery packs batt1 and batt4 fails. The crosslink failure may result in battery batt1 (indicated by battery index 1802), which has only half of its available energy, having HV channel 1ch1 (of the failed crosslink) with only one-quarter of the available energy of either bus 2 or bus 3. In addition, the crosslink failure may result in battery batt4 (indicated by battery index 1802), which has the same available energy as batteries batt2, batt3, batt5, and batt6, having HV channel 4ch4 (of the failed crosslink) with only half of the available energy of either bus 2 or bus 3 (indicated by HV bus index 1804). As shown in engine index 1806, the energy optimization function may adjust the reference command by using the estimated remaining discharge time per bus / channel. For example, the energy optimization function may reduce the torque / speed / usage associated with engines EE1 and EE12 by adjusting the reference command associated with engines EE1 and EE12 (powered only by battery batt1 via HV channel ch1) to a lower value (e.g., "power reference" is approximately 0.6), and increase the torque / speed / usage associated with engines EE2-EE11 by adjusting the reference command associated with engines EE2-EE11 to a higher value (e.g., "power reference" is approximately 1.2). As shown in HV channel index 1808, the energy optimization function may reduce the reference command associated with HV channel ch1 (of the faulty crosslink) connected to battery batt1, and increase the respective reference commands associated with HV channel ch4 and HV buses bus2 and bus3 to prevent HV channel ch1 from running out of energy before channel ch4 and buses bus2 and bus3.

[0154] In some embodiments, the flight control system may include one or more engine thermal management functions. Engine thermal imbalance can have several causes, including uneven propeller commands, differences in heat dissipation, differences in battery and / or engine health, and battery or engine failure. Electric engines may have thermal operating limits, and the engine thermal management function may be configured to prevent the engine from experiencing unacceptable operating characteristics (e.g., operating beyond a risk threshold) and exceeding operating limits, including failure conditions.

[0155] In some embodiments, at least one (e.g., each) electric engine (e.g., at least one processor associated with each engine (e.g., part of it, and / or connected to it)) may be configured to estimate the remaining time of the engine at the current power setting (e.g., as part of system 1100 and / or method 1200). In some embodiments, at least one electric engine may include an engine time estimator (e.g., using the electric engine's inverter or processor) configured to determine the temperature of one or more engine components. For example, various engine sensors may collect temperature data associated with one or more engine components, such as a control board, DC link capacitor, rapid discharge, motor control unit (MCU), power module junction, stator windings, oil, or rotor magnet of an EPU. In some embodiments, the engine time estimator may normalize the temperature in terms of time. In some embodiments, the engine time estimator may normalize the determined (e.g., calculated, measured, or received from) temperatures of different components (e.g., rotor magnets, oil, control board, DC link capacitor, motor control unit (MCU), rapid discharge, power module junction, stator windings, etc.) having different time constants and temperature limits, using a time-based metric for temperature. For example, different components of an electric engine may have different temperature limits (e.g., minimum and maximum temperature limits). In some embodiments, the engine time estimator may estimate the remaining time based on the component having the lowest estimated remaining time. For example, the engine time estimator may communicate the lowest estimated remaining time to the FCS as the estimated remaining time for the engine at the current power setting.

[0156] Figure 19 illustrates exemplary temperature-versus-time plots of components for estimating remaining time (for example, used as part of system 1100 and / or method 1200) according to various embodiments. The engine time estimator may generate plots of one or more components for summarizing the temperatures of different components with different temperature limits into time-based metrics that enable visualization and / or data structures (e.g., usable by engine thermal management functions) of the minimum time the engine overheats. In some embodiments, the engine time estimator may consider all components of the engine. In some embodiments, the engine time estimator may consider a subset of components. In some embodiments, each phase of flight may have a different power output.

[0157] Figure 20 illustrates exemplary gauges for displaying information related to estimated remaining time (e.g., used as part of system 1100 and / or method 1200) according to various embodiments. In some embodiments, the gauges may display the proximity of torque or other aircraft parameters to defined limits. While the gauges in Figure 20 are shown in a particular combination of layout and visual elements, it should be understood that other modifications may be used to indicate the proximity of the current operational state of the aircraft to limits and / or remaining time quantities in the operational state of the aircraft (e.g., safety time, the remaining time until the engine reaches or exceeds the limit). In some embodiments, the information displayed on the gauges may be based on information derived according to features described with respect to Figure 19 or one or more of Figures 21-25. A gauge (e.g., gauges 2002, 2004, 2006) may include a torque arc (e.g., torque arc 2008) showing the available torque from the engine in operation, an indicator (e.g., indicator 2010) showing the torque in use, limit markers (e.g., continuous limit marker 2012 (CONT), takeoff limit marker 2014 (TO), emergency limit marker 2016 (EMRG)) showing the currently active torque limit, a red line (e.g., red line 2018) showing the maximum torque, and a boost bar (e.g., boost bar 2020) showing the time remaining at the current rated maximum torque. For example, gauge 2002 displays a continuous limit marker 2012 to indicate that the currently active torque limit is a continuous limit, which is associated with an infinite remaining time, as shown in boost bar 2022. Gauge 2004 displays a takeoff limit marker 2014 to indicate that the currently active torque limit is a takeoff limit, which is associated with the remaining time being below a first predetermined threshold (e.g., 243 seconds below 300 seconds), as shown on the boost bar 2020. Gauge 2006 displays an emergency limit marker 2016 to indicate that the currently active torque limit is an emergency limit, which is associated with the remaining time being below a second predetermined threshold (e.g., 27 seconds below 30 seconds), as shown on the boost bar 2024.In some embodiments, the second predetermined threshold includes a time less than the first predetermined threshold.

[0158] In some embodiments, the engine time estimator may be configured to determine one or more temperature parameters. For example, the engine time estimator may be configured to collect inputs from various sources (e.g., sensors, estimation algorithms, etc.). In some embodiments, at least one of the engine time estimator or the flight control system may be configured to identify the validity of the collected inputs. For example, at least one of the flight control system or the engine time estimator may be configured to validate one or more communication signals (e.g., fault checks, loss of communication within a certain timeout window, etc.). In some embodiments, the engine time estimator may be configured to calculate an integrated temperature value based on the input validity.

[0159] In some embodiments, the engine time estimator may be configured to detect one or more temperature disturbances (e.g., when the temperature reaches the FCS warning threshold, when the temperature reaches the engine action level, thermal disturbance on the power stage, low oil level, low oil flow, no inverter oil flow, no stator oil flow, no gearbox oil flow, cutoff of airflow to the HEX, temperature sensor degradation, loss of temperature sensing, single-phase short circuit, etc.). For example, the engine time estimator may be configured to perform a validity check of the integrated temperature value. In some embodiments, detecting one or more temperature disturbances may include determining whether the integrated temperature value is outside the rated range. In some embodiments, detecting one or more temperature disturbances may include determining whether the integrated temperature value is outside the limit range. In some embodiments, detecting one or more temperature disturbances may include determining whether the engine is thermally degrading (e.g., excessive ΔT increase due to oil leak).

[0160] In some embodiments, the engine time estimator may run a background thermal model. For example, the background thermal model may predict the temperature of at least one (e.g., each) component of the engine (e.g., rotor magnets, oil, control board, DC link capacitor, motor control unit (MCU), rapid discharge, power module junction, stator windings, etc.). In some embodiments, the background thermal model may utilize multiple nodes to predict the temperature of at least one component of the engine. Additionally or alternatively, the background thermal model may utilize a detailed loss function to predict the temperature of at least one major component of the engine. In some embodiments, the background thermal model may be fine-tuned and / or pass multiple rounds of testing to improve the accuracy of temperature predictions.

[0161] In some embodiments, the engine time estimator may run a time-out model. For example, the time-out model may utilize information from a background thermal model to determine the current thermal state (e.g., current temperature) of at least one engine component. In some embodiments, the time-out model may perform forward predictions to determine (e.g., predict) the time remaining until one or more thermal limits are reached for at least one (e.g., each) component of the engine, based on the determined current thermal state of the component. In some embodiments, the time-out model may take external ambient conditions into account. In some embodiments, the time-out model may include a reduced-order thermal model with a simplified loss function. In some embodiments, the time-out model may include an analytical fit exponential rise curve with flattening for at least one (e.g., each) engine component of interest (e.g., rotor magnets, oil, control board, DC link capacitor, motor control unit (MCU), rapid discharge, power module junction, stator winding, etc.). In some embodiments, the remaining time model may include a machine learning model (e.g., a neural network model, as well as a model trained on engine temperature values, flight parameters, and / or energy states at one or more flight phases) trained to accept one or more current thermal states of one or more engine components of the engine as input to output the remaining time associated with the engine. In some embodiments, the engine time estimator may generate a machine learning model based on training data (e.g., data from simulations, historical data, and background thermal models).

[0162] In some embodiments, at least one of the engine thermal management function or the flight control system may be configured to perform one or more protective actions against overheating. In some embodiments, the protective action may include one or more of the following: transmitting a warning (e.g., to the flight control system, the engine, the aircraft pilot, or the nearest landing area via GPS), performing a torque ramp-down (e.g., reducing the torque required by at least one EPU over time and / or adjusting at least one command), or performing an emergency shutdown.

[0163] In some embodiments, at least one (e.g., each or more) engine may communicate its respective estimated remaining time to the flight control system.

[0164] Figure 21 illustrates exemplary temperature thresholds (e.g., available in System 2200, Method 2300, and / or Method 2400) for various embodiments. Threshold 2131 may be the maximum rated point / soft sealing, and when engine parameters exceed threshold 2131, the engine may send a warning to the FCC. In response to receiving the warning, the FCC may take one or more protective actions. For example, the FCC may issue a warning to the pilot, instruct a reduction in engine torque output, and / or instruct the engine to shut down. Exceeding threshold 2132 may indicate that the engine is approaching its operating limit, and the flight control system may automatically reduce engine torque output in response to exceeding this threshold to prevent the engine from overheating. Threshold 2133 may be the operating limit, and the flight control system may take one or more emergency actions in response to reaching this threshold. For example, the flight control system may shut down the engine in question. In some embodiments, commands received from the pilot (e.g., one or both pilot inceptors, buttons, switches) may be configured to override one or more emergency actions performed by the flight control system in order to exceed an operational limit threshold.

[0165] Figure 22 is a functional block diagram of an exemplary system 2200 for a VTOL aircraft, including a torque limit calculation function (e.g., torque limit calculation function 2226), consistent with the disclosed embodiments. It is understood that a particular machine (e.g., an aircraft) may use the exemplary system 2200 to implement improvements in the technical fields of aircraft safety, stability, reliability, and efficiency. For example, some embodiments may involve a dynamically changing torque command based on at least one aircraft condition, which can be used by the propulsion unit to help make the aircraft safer, more stable, easier to fly, more reliable, and more efficient, for example, during different flight phases or modes, consistent with the disclosed embodiments. As illustrated in Figure 22, the system 2200 may include a torque limit calculation function 2210, a control allocation function 2220, and a DCPS 2230. The system 2200 may be implemented by a microprocessor-based controller that executes software code stored in a storage medium to implement the functions described herein. System 2200 can also be implemented in hardware or as a combination of hardware and software. System 2200 can be configured to repeatedly perform a single step or sequence until a desired or commanded result is obtained. Please understand that many of the conventional functions of the control system are not shown in Figure 22 for the sake of clarity.

[0166] In some embodiments, the torque limit calculation function 2210 may be configured to receive at least one of the following as inputs: system faults, pilot inputs, or engine data. For example, the torque limit calculation function 2210 may be configured to receive one or more of the following (e.g., from multiple engines of an aircraft or from each engine): scheduled propeller torque limits 2201, engine modes 2202, engine fault status 2203, continuous engine torque limits 2204, takeoff torque limits 2205, emergency engine torque limits 2206, overheat torque limits 2207, predicted duration at maximum takeoff torque 2208, or predicted duration at maximum emergency torque 2209. In some embodiments, one or more inputs may be preset based on at least one of the following: aircraft design specifications, structural limitations, or cooling capacity. For example, one or more inputs may be associated with regulatory values.

[0167] In some embodiments, the system 2200 may include an engine power rating management function 2215. The engine power rating management function 2215 may be configured to select an appropriate torque rating (e.g., an engine continuous torque limit 2204, an engine takeoff torque limit 2205, an engine emergency torque limit 2206) for at least one engine (e.g., each or more engines of the aircraft) based on the flight phase (e.g., hovering, cruising / continuous, transition, takeoff / landing) and one or more operational requirements. For example, function 2215 may be configured to select the engine takeoff torque limit 2205 as the torque rating based on the determination that the aircraft is in the takeoff or landing flight phase. Additionally or alternatively, function 2215 may be configured to select the engine continuous torque limit 2204 as the torque rating based on the determination that the aircraft is in the cruising flight phase. Additionally or alternatively, function 2215 may be configured to select the engine emergency torque limit 2206 as the torque rating based on the detection of one or more system failures (e.g., engine failure, damaged aircraft component, etc.). In some embodiments, the power rating management function 2215 may include an automatic function configured to dynamically switch between torque ratings based on flight phase, operational requirements, and vehicle dynamics (e.g., aircraft conditions, system failure). In some embodiments, the engine power rating management function 2215 may include a switch configured to receive pilot input via a toggle function.

[0168] In some embodiments, the system 2200 may include an engine thermal management function 2221. The engine thermal management function 2221 may be configured to manage (e.g., adjust, configure, set, and / or monitor) engine temperatures to ensure that (e.g., each, at least one) engines operate within safe parameters to prevent overheating and damage to engine components. For example, the engine thermal management function 2221 may be configured to determine an engine rated torque limit 2216 for at least one engine based on the maximum output of at least one aircraft engine output from an engine power rating management function 2215 (e.g., for each engine or a group of aircraft engines).

[0169] The engine rated torque limit 2216 may correspond to the maximum torque that the engine may be able to generate under normal continuous operating conditions. In some embodiments, from at least one (e.g., each) engine, the engine thermal management function 2221 may receive one or more of the following: the engine overheat torque limit 2207, the predicted duration at maximum takeoff torque 2208, or the predicted duration at maximum emergency torque 2209. Based on at least one of the predicted duration at maximum takeoff torque 2208 or the predicted duration at maximum emergency torque 2209, the engine thermal management function 2221 may determine the limits to use to determine the steady-state engine torque limit 2222 (e.g., the engine continuous torque limit 2204, the engine takeoff torque limit 2205, the engine emergency torque limit 2206, the engine overheat torque limit 2207, and the engine rated torque limit 2216). For example, based on the determined amount of remaining time for takeoff torque, which may be represented by a timer for the remaining time (e.g., greater than 0, the aircraft can maintain its operation at maximum takeoff torque) in 2217 and the predicted duration 2208 at maximum takeoff torque indicated by the remaining time in 2218, the engine thermal management function 2221 may use the engine rated torque limit 2216 to determine the steady-state engine torque limit 2222. Additionally or alternatively, based on the determination that there is no remaining time for takeoff torque, which may be represented by a timer for the remaining time (e.g., less than or equal to 0) in 2217 and the predicted duration 2208 at maximum takeoff torque indicated by the remaining time in 2218, the engine thermal management function 2221 may use the engine continuous torque limit 2204 to determine the steady-state engine torque limit 2222. Additionally or alternatively, based on the fact that there is no time remaining for the emergency torque, the engine thermal management function 2221 may use the engine overheat torque limit 2207 to determine the steady-state engine torque limit 2222, which may be represented by a timer for the predicted duration 2209 at maximum emergency torque, indicating that there is no time remaining at 2218 (e.g., less than or equal to 0, the aircraft cannot maintain operation at maximum emergency torque).In some embodiments, the engine heat management function 2221 may be configured such that a time buffer exists between timer 2217 and timer 2218 (for example, so that timer 2217 and timer 2218 do not reach zero at the same time).

[0170] In some embodiments, the engine thermal management function 2221 may be configured to return the torque limit to the normal torque rating after receiving an instruction for engine cooling (e.g., based on and in response to such instruction). For example, after setting the engine overheat torque limit 2207 as the torque limit for the engine, the engine thermal management function 2221 may receive an instruction that the engine temperature has fallen below a predetermined temperature threshold. Additionally or alternatively, the engine thermal management function 2221 may receive an instruction that the remaining time for the predicted duration at the torque rating (e.g., timer 2218) is greater than a predetermined time threshold (e.g., indicating sufficient engine cooling). Based on the instruction, the engine thermal management function 2221 may be configured to set one of the torque limits 2204, 2205, or 2206 as the torque limit for the engine.

[0171] In some embodiments, the torque limit calculation function 2210 may be configured to determine one or more steady-state engine torque limits 2222. Determining one or more steady-state engine torque limits 2222 may include at least one of receiving, retrieving, or checking one or more scheduled propeller torque limits. In some embodiments, one or more scheduled propeller torque limits 2201 may be determined based on at least one of one or more RPM limits and / or structural load limits. For example, a structural load limit may be associated with loads imparted by one or more of the engine on the propeller hub, propeller rotation, propeller blades, mounting between the propeller and the engine, mounting between the engine and the boom, mounting between the boom and the wing, or the boom itself. Additionally or alternatively, one or more scheduled propeller torque limits 2201 may include one or more predetermined values ​​specified for a particular aircraft configuration. For example, one or more scheduled propeller torque limits 2201 may be specified based on the aircraft's propeller design or structural limitations. In some embodiments, determining one or more steady-state engine torque limits 2222 may include determining engine availability 2211. For example, the engine availability determination 2211 may include determining whether at least one (e.g., each) engine is currently operating and available for control, based on one or more of the engine modes 2202 (e.g., inverter internal state machine state, a coefficient of 0 in standby / listening mode, a coefficient of 0 in storage mode, a coefficient of 1 in closed-loop torque command mode, etc.) or engine failure statuses 2203. In some embodiments, the engine availability determination 2211 may output an engine availability coefficient 2212, which may indicate the level of functionality (e.g., relative to expected functionality or full functionality) associated with one or more engines.For example, the engine availability factor 2212 may include values ​​such as 0 for not operating, 1 for being fully operational, and 0.5 for indicating a failure condition (e.g., losing one of two winding sets, causing the engine to produce only about half the normal amount of torque). In some embodiments, based on torque limits determined by the engine thermal management function 2221 (e.g., continuous engine torque limit 2204, rated engine torque limit 2216, overheated engine torque limit 2207), the torque limit calculation function 2210 may, in 2213, combine (e.g., multiply) the engine availability factor 2212 with the torque limits. In 2214, the torque limit calculation function 2210 may compare the scheduled propeller torque limit 2201 with the output of 2213 and output the lower limit as the steady-state engine torque limit 2222.

[0172] In some embodiments, the torque limit calculation function 2210 may be configured to determine one or more dynamic engine torque limits 2224. Determining one or more dynamic engine torque limits 2224 may include, in 2219, combining (e.g., multiplying) the engine availability coefficient 2212 output from the engine availability determination 2211 with the engine emergency torque limit 2206. In some embodiments, the steady-state engine torque limit 2222 for the engine may include a limit lower than the dynamic engine torque limit 2224 for the engine.

[0173] The control allocation function 2220 may be configured to determine one or more allocated torque commands 2223. In some embodiments, the one or more allocated torque commands 2223 may include one or more propeller torque commands. In some embodiments, the control allocation function 2220 may be configured to determine one or more allocated torque commands 2223 based on at least one of the steady-state engine torque limit 2222, the predicted duration at maximum takeoff torque 2208, or the predicted duration at maximum emergency torque 2209.

[0174] The DCPS2230 may be configured to determine one or more engine torque commands 2232. In some embodiments, the one or more engine torque commands 2232 may include one or more modified propeller torque commands. In some embodiments, the DCPS2230 may be configured to determine one or more engine torque commands 2232 based on at least one of an assigned torque command 2223 or a dynamic engine torque limit 2224.

[0175] Figure 23 is a flowchart illustrating exemplary methods 2300 for engine thermal management of a particular machine (e.g., an aircraft) according to some embodiments of the present disclosure, thereby improving aircraft technology (e.g., aircraft safety, thermal efficiency, payload capacity, structural integrity). The steps of Method 2300 may be implemented, for example, by System 1000 of Figure 10, on any aircraft component of Figures 1-8, 9A-9E, 10-12, 13A-13D, 14-22, 24, 25, or on any flight control computer (e.g., a method performed by a computer) or flight control system, or by using them in other ways. For example, an aircraft flight control computer may be configured to perform one or more steps of Method 2300.

[0176] The steps of Method 2300 can be activated or modified as needed, based on at least one aircraft flight condition, at least one maneuver, and / or at least one operating requirement. It is understood that the illustrated Method 2300 can be modified by changing the order of the steps and by including additional steps. It is also understood that the complexity of Method 2300 would be impossible, or at least extremely impractical, to implement effectively by a human user, especially considering that these functionalities are implemented (e.g., in real time) while the energy used by the aircraft and the battery and engine conditions are consistently changing while the aircraft is in flight (including takeoff or landing). Furthermore, engine thermal management (e.g., using Method 2300) improves the safety and performance capabilities of the aircraft (e.g., how close the engine can get to maximum performance without exceeding thermal limits), as well as the lifespan of different components (e.g., batteries, engines), without requiring the direct involvement of a user (e.g., pilot). In some embodiments, the aircraft of system 1000 may comprise an electric aircraft comprising multiple effectors, including multiple electric propulsion units (EPUs) and multiple battery packs that supply power to the multiple electric propulsion units.

[0177] In step 2302, system 1000 may determine one or more desired commands for the electric aircraft. For example, one or more desired commands may be one or more force or moment commands (e.g., inputs to control allocation 1029 in Figure 10), a desired position and / or turn rate command (e.g., output of turn rate command model 1004 in Figure 10), a desired position and / or lateral velocity command (e.g., output of lateral velocity command model 1006 in Figure 10), one or more desired altitude, vertical velocity, or vertical acceleration commands (e.g., output of climb command model 1008 in Figure 10), one or more desired position, longitudinal velocity, or longitudinal acceleration commands (e.g., output of forward velocity command model 1010 in Figure 10), alternative commands to achieve a desired change (e.g., based on feedback as considered in Figure 10), one or more desired total The commands may include at least one of the following commands associated with a desired movement of the aircraft (e.g., desired by the aircraft's pilot, desired by the flight control system, desired by the autopilot system): force (e.g., calculated based on the outputs of feedback 1012, 1012, 1016, 1018, 1022 and feedforward 1014, 1020 in Figure 10), one or more moment commands (e.g., determined by the inner-loop control law 1028 in Figure 10), linear commands, angular acceleration commands, or signals (e.g., pilot inputs from pilot input devices, inputs from the autopilot system, signals generated by the flight control system or computer).

[0178] In step 2304, system 1000 may receive engine information relating to at least one (e.g., each) electric propulsion unit (EPU) among a plurality of EPUs. For example, system 1000 may also monitor the engine status of at least one (e.g., each) propulsion unit among a plurality of electric propulsion units. For example, system 1000 may receive engine information (e.g., engine status) associated with the EPU from at least one (e.g., each) EPU among a plurality of EPUs (e.g., via a digital communication interface, via the EPU's inverter and / or processor). Engine information may include one or more of the following: the temperature of at least one (e.g., each) engine (e.g., current temperature, predicted temperature, temperature against limits); one or more limits associated with at least one engine (e.g., temperature limits); time associated with the temperature of at least one (e.g., each) engine (e.g., time remaining under current rating such as continuous, takeoff, or emergency, time remaining until limits (e.g., thermal limits) are reached); or status (e.g., fault conditions, currently operating (e.g., active) and controllable engines(s), currently inoperable and uncontrollable engines(s)). In some embodiments, engine information may include one or more of the following: engine overheat torque limits, predicted duration at maximum takeoff torque, or predicted duration at maximum emergency torque.

[0179] In step 2306, the system 1000 may generate control commands for multiple effectors based on the received engine information. For example, the system 1000 may generate control commands to manage (e.g., adjust, configure, set, and / or monitor) engine temperature to ensure that at least one (e.g., each) engine operates within safe parameters and prevents overheating and damage to engine components. In some embodiments, the system 1000 may generate control commands to limit the actions performed by at least one engine or other component of the aircraft in order to reduce strain or risk to the relevant components, while still controlling the effector based on the desired command (e.g., having minimal impact on a reference state, flight envelope, etc.). In some embodiments, the system 1000 may determine an engine rated torque limit for at least one (e.g., each) engine. The engine rated limit may include the maximum torque that the engine may be able to generate under normal (e.g., continuous) operating conditions. In some embodiments, the system 1000 may determine the engine rated torque limit based on the received engine information. A detailed explanation of the generation of control commands (e.g., assigned torque command 2223 or torque command 2232) is provided in the description of Figure 22 above. In some embodiments, generating control commands for multiple effectors may be based on one or more aircraft conditions. For example, system 1000 may generate control commands based on the current aircraft conditions.

[0180] In some embodiments, generating control commands may be based further on the energy status of multiple battery packs. For example, system 1000 may monitor the energy status of multiple battery packs. In some embodiments, system 1000 may be communicatively coupled to one or more battery management systems ("BMS") of an aircraft (e.g., via physical connections such as bus and / or channel systems, or via a digital communication interface). In some embodiments, an aircraft may include a single BMS configured to manage all battery packs on the aircraft. In some embodiments, at least one (e.g., each) battery pack may include its own BMS. In some embodiments, system 1000 may receive battery information (e.g., energy status) associated with at least one (e.g., each) battery from one or more BMS associated with multiple battery packs. Battery information may include one or more of the following: available energy, remaining discharge time (e.g., for at least one crosslink, at least one independent high-voltage bus, at least one high-voltage channel), energy state (SOE), charge state (SOC), power state (SOP), health state (SOH), fault conditions (e.g., short circuit or overcurrent, whether the battery is active / functioning), or temperature state (SOT) of at least one battery pack.

[0181] In step 2308, system 1000 may control a plurality of effectors in accordance with the generated control commands to satisfy (e.g., respond to, satisfy, deal with, and based on) a desired command of an electric aircraft. For example, system 1000 may transmit at least one (e.g., each) generated control command to its respective effector, and the effector may move (e.g., in accordance with) the controlled command.

[0182] In some embodiments, Method 2300 can result in optimized engine performance by maximizing the performance of at least one (e.g., each) engine while remaining within the engine temperature limits. In effect, Method 2300 can result in maximizing the available range and can also enhance safety by avoiding a dangerous situation in which one or more engines fail due to thermal overload before the others and are no longer able to generate thrust.

[0183] Figure 24 is a block diagram illustrating exemplary methods 2400 for engine thermal management and energy optimization of a particular machine (e.g., an aircraft) according to various embodiments of the present disclosure, thereby improving aircraft technology (e.g., aircraft safety, thermal efficiency, payload capacity, structural integrity). The steps of Method 2400 may be implemented, for example, by the system 1000 of Figure 10 being performed on any aircraft component of Figures 1-8, 9A-9E, 10-12, 13A-13D, 14-23, and 25, or on any flight control computer (e.g., a method performed by a computer) or flight control system, or by using them in other ways. For example, an aircraft flight control computer may be configured to perform one or more steps of Method 2400. Furthermore, the complexity of Method 2400 is understood to be such that, in particular, while the energy and battery conditions used by the aircraft are consistently changing during flight (including takeoff or landing), these functionalities must be implemented (e.g., in real time), making effective implementation by a human user impossible or at least highly impractical. Moreover, energy optimization (e.g., using Method 1200) improves the safety and range capability of the aircraft (e.g., how much energy is available to power the aircraft) as well as the lifespan of different components (e.g., batteries) without requiring the direct involvement of a user (e.g., a pilot). Figure 24 illustrates an example of energy optimization and thermal management associated with torque and power, but it should be further understood that Figure 24 can be associated, alternatively or additionally, with any parameters related to aircraft effectors such as speed and / or current. The steps of Method 2400 can be activated or adjusted as needed based on the aircraft's flight conditions, maneuvering, or operating requirements. It will be understood that the illustrated method 2400 can be modified by changing the order of the steps and by including additional steps.

[0184] In step 2401, the thermal management function may calculate at least one torque or power adjustment coefficient using at least one remaining time until at least one thermal limit is reached. In some embodiments, the thermal management function may be configured to receive one or more of the following: battery information, electrical wiring interconnect system (EWIS) information, or engine information. In some embodiments, the thermal management function may receive at least one (e.g., each) remaining time from at least one (e.g., each, multiple) engine. In some embodiments, the thermal management function may calculate at least one torque or power adjustment coefficient based on one or more of the received battery information, EWIS information, or engine information.

[0185] In step 2402, the energy optimization function may calculate a power adjustment coefficient, for example, using the difference in available energy (AE) per bus. In some embodiments, the energy optimization function may receive at least one crosslink status of a plurality of crosslink statuses. In some embodiments, the energy optimization function may receive the difference in available energy per bus from at least one BMS.

[0186] In step 2403, the flight control system may be configured to adjust at least one electric engine reference value (e.g., reference command) based on values ​​calculated by thermal management and energy optimization functions. In some embodiments, the flight control system may determine a new reference state to which control allocation should be resolved. In some embodiments, the flight control system may calculate a deviation (e.g., difference) of the current state (e.g., configuration, effector position, effector output) from the new reference state (e.g., associated with at least one electric engine reference value, reference torque setting).

[0187] In 2404, the flight control system may be configured to recalculate force / moment commands. A detailed explanation of force / moment command recalculation is provided in the description of Figure 11 above.

[0188] Figures 25A and 25B illustrate exemplary scenarios in which both thermal management and energy optimization are carried out according to several embodiments (e.g., using and / or following System 1100, Method 1200, Method 2300, and / or Method 2400). As shown in Chart 2500a of Figure 25A, all electric engines have the same initial torque reference value (e.g., reference command, τ ref Starting at ), HV bus 2 has more energy than buses 1 and 3 (indicated by B2 and B5 being fuller than B1 and B4 as well as B3 and B6), and electric engine (EE) 8 is hotter than the other engines. As shown in Chart 2500b of Figure 25B, thermal management and energy optimization are carried out by adjusting the initial torque baselines. For example, the baselines associated with electric engines 2, 5, and 11 in bus 2, which has the most available energy, are increased (e.g., to increase the use of electric engines 2, 5, and 11), and the baseline associated with electric engine 8 is decreased (e.g., to decrease its use compared to the other engines). In addition, the baselines associated with the electric engines in buses 1 and 3 are adjusted so that the use of the engines associated with each of buses 1 and 3 is lower than the use of engines 2, 5, and 11. Implementing both thermal management and energy optimization can result in full energy utilization in all HV buses while staying within the engine thermal limits.

[0189] In some embodiments, engine thermal management may take precedence over energy optimization. Alternatively, energy optimization may take precedence over engine thermal management.

[0190] Additional aspects of this disclosure may be further described through the following clauses. 1. A method that is performed on a computer, To determine one or more desired directives for an aircraft, Based on the one or more desired directives and one or more aircraft conditions, determine at least one reference directive, Monitoring the energy state of multiple battery packs of the aircraft, wherein at least one first battery pack among the multiple battery packs is electrically isolated from at least one second battery pack among the multiple battery packs. Adjusting the at least one reference command based on the monitored energy state of the plurality of battery packs, To generate control commands for multiple effectors of the aircraft based on the adjusted reference command, A method performed by a computer, comprising controlling the plurality of effectors according to the generated control commands in order to satisfy the one or more desired commands of the aircraft. 2. The method by which the monitored energy state is performed on the computer described in Clause 1, including the available energy of the plurality of battery packs. 3. A method by which the monitored energy state is performed by a computer as described in Clause 1 or 2, including the remaining discharge time associated with one or more high-voltage channels. 4. A method performed on a computer according to any one of the clauses 1 to 3, wherein the monitored energy state includes the difference in energy states between at least two of the plurality of battery packs. 5. A method by which the monitored energy state is performed by a computer as described in any one of Clauses 1 to 4, including the difference in power consumption between at least the first and second engines of the multiple electric propulsion units of the aircraft. 6. The method, performed by a computer as described in Clause 5, wherein the first engine has lower power consumption than the second engine, and the generated control commands cause a reduction in power consumption from the second engine. 7. A method by which the adjustment of the at least one reference directive is performed by a computer as described in any one of the clauses 1 to 6, further based on the state of one or more engines of the aircraft's multiple electric propulsion units. 8. A method performed by a computer as described in any one of the clauses 1 to 7, wherein adjusting the at least one reference instruction includes updating one or more reference instructions to optimize the remaining discharge time across one or more high-voltage channels. 9. A method by which the at least one reference directive is executed by a computer as described in any one of Clauses 1 to 8, wherein the reference directive includes one or more of the reference engine directive, reference power directive, reference torque directive, or reference speed directive. 10. A method by which the one or more aircraft conditions are performed on a computer as described in any one of Clauses 1 to 9, including one or more of the vehicle dynamics, flight conditions, or the status of at least one aircraft component. 11. A method performed on a computer controlling an aircraft, comprising a plurality of effectors including a plurality of electric propulsion units and a plurality of battery packs that supply power to the plurality of electric propulsion units, To determine one or more desired instructions for the aircraft, Based on the aforementioned desired directives and one or more aircraft conditions, determine at least one reference directive, Monitoring the energy state of the plurality of battery packs, wherein at least one first battery pack among the plurality of battery packs is electrically isolated from at least one second battery pack among the plurality of battery packs. Adjusting the at least one reference command based on the monitored energy state of the plurality of battery packs, Based on the adjusted reference command, generate control commands for the plurality of effectors, A method performed by a computer, comprising controlling the plurality of effectors according to the generated control commands in order to satisfy the one or more desired commands of the aircraft. 12. Flight control computer, One or more memory devices that store processor-executable instructions, A flight control computer comprising: one or more processors configured to execute the aforementioned instructions and cause the flight control computer to perform the methods performed by the computer described in any one of the clauses 1 to 11. 13. A non-temporary computer-readable medium storing computer-readable instructions, wherein when the computer-readable instructions are executed by at least one processor, the at least one processor causes the at least one processor to execute the computer-executed method described in any one of the clauses 1 to 11. 14. An aircraft having at least one flight control computer configured to perform the computer-operated methods described in any one of the clauses 1 to 11. 15. A flight control system, At least one memory location that stores instructions, A flight control system comprising: at least one processor configured to execute the aforementioned instructions and carry out the methods performed on a computer as described in any one of the clauses 1 to 11. 16. An aircraft flight control system, At least one memory location that stores instructions, The system comprises at least one processor configured to execute the aforementioned instructions and perform one or more operations, wherein the operations are To determine one or more desired directives for an aircraft, Based on the one or more desired directives and one or more aircraft conditions, determine at least one reference directive, Monitoring the energy state of multiple battery packs of the aircraft, wherein at least one first battery pack among the multiple battery packs is electrically isolated from at least one second battery pack among the multiple battery packs. Adjusting the at least one reference command based on the monitored energy state of the plurality of battery packs, To generate control commands for multiple effectors of the aircraft based on the adjusted reference command, A flight control system comprising controlling the plurality of effectors according to the generated control commands to satisfy the one or more desired commands of the aircraft. 17. The flight control system according to Clause 16, wherein the monitored energy state includes the available energy of the plurality of battery packs. 18. The flight control system according to Clause 16 or 17, wherein the monitored energy state includes the remaining discharge time associated with one or more high-voltage channels. 19. The flight control system according to any one of the clauses 16 to 18, wherein the monitored energy state includes the difference in energy states between at least two of the plurality of battery packs. 20. A flight control system according to any one of the clauses 16 to 19, wherein the monitored energy state includes the difference in power consumption between at least a first engine and a second engine of a plurality of electric propulsion units of the aircraft. 21. The flight control system according to Clause 20, wherein the first engine has lower power consumption than the second engine, and the generated control commands cause a reduction in power consumption from the second engine. 22. A flight control system according to any one of the clauses 16 to 21, wherein adjusting the at least one reference directive further depends on the state of one or more engines of the aircraft's multiple electric propulsion units. 23. A flight control system according to any one of the clauses 16 to 22, wherein adjusting the at least one reference command includes updating one or more reference commands to optimize the remaining discharge time across one or more high-voltage channels. 24. A flight control system according to any one of clauses 16 to 23, wherein the at least one reference command includes one or more of the following: a reference engine command, a reference power command, a reference torque command, or a reference speed command. 25. A flight control system according to any one of the clauses 16 to 24, wherein the one or more aircraft conditions include one or more of vehicle dynamics, flight conditions, or the status of at least one aircraft component. 26. A non-temporary computer-readable medium storing one or more instructions, wherein when one or more instructions are executed by at least one processor, the medium causes the at least one processor to perform an operation, and the operation is To determine one or more desired directives for an aircraft, Based on the one or more desired directives and one or more aircraft conditions, determine at least one reference directive, Monitoring the energy state of multiple battery packs of the aircraft, wherein at least one first battery pack among the multiple battery packs is electrically isolated from at least one second battery pack among the multiple battery packs. Adjusting the at least one reference command based on the monitored energy state of the plurality of battery packs, To generate control commands for multiple effectors of the aircraft based on the adjusted reference command, A non-temporary computer-readable medium comprising controlling the plurality of effectors according to the generated control commands to satisfy the one or more desired commands of the aircraft. 27. A computer-readable medium storing computer-readable instructions, wherein when the computer-readable instructions are executed by at least one processor, the computer-readable medium causes at least one processor to execute the method described in any one of the clauses 1 to 11. 28. A method that is performed on a computer, To determine one or more desired directives for an aircraft, Receiving engine information relating to at least one of the multiple electric propulsion units (EPUs) of the aircraft, wherein the engine information includes at least one temperature associated with the at least one EPU. Based on the received engine information, control commands are generated for multiple effectors of the aircraft. A method performed by a computer, comprising controlling the plurality of effectors according to the generated control commands in order to satisfy the one or more desired commands of the aircraft. 29. A method of performing the engine information in a computer as described in Clause 28, further including a time related to at least one temperature generated by the at least one EPU. 30. The method performed by the computer described in Clause 29, wherein the time is generated by determining the current temperature of each component of a plurality of engine components for at least one EPU. 31. The method performed by a computer according to Clause 30, wherein the time is generated by normalizing the determined temperatures of the plurality of engine components, including one or more engine components having different temperature limits, using a time-based metric for temperature. 32. The method performed by the computer described in Clause 30, wherein the time is generated based on a prediction of the remaining time at the current torque rating associated with the engine of the EPU. 33. A method by which the engine information is performed on a computer as described in any one of the clauses 28 to 32, further including the status of at least one engine associated with the plurality of EPUs. 34. A method of execution on the computer described in Clause 33, wherein the status includes one of active or inactive. 35. A method by which the generation of the control commands is performed by a computer as described in any one of Clauses 28 to 34, further based on the maximum torque that the engine can generate under normal operating conditions. 36. A method by which the generation of the control commands is performed by a computer as described in any one of the clauses 28 to 35, further based on one or more aircraft conditions. 37. The method by which the one or more aircraft conditions are performed on a computer as described in Clause 36, including one or more of the vehicle dynamics, flight conditions, or the status of at least one aircraft component. 38. The method of execution on the computer described in Clause 29, wherein the time includes the remaining time determined by the machine learning model of the engine. 39. A method by which the engine information is performed in a computer according to any one of the clauses 28 to 38, further including the temperature associated with the engine of the at least one EPU. 40. A method for controlling an aircraft, comprising a plurality of effectors, each comprising a plurality of electric propulsion units (EPUs) and a plurality of battery packs that supply power to the plurality of electric propulsion units, To determine one or more desired instructions for the aircraft, Receiving engine information relating to at least one of the plurality of EPUs, wherein the engine information includes at least one temperature associated with the at least one EPU. Based on the received engine information, control commands for the plurality of effectors are generated, A method comprising controlling the plurality of effectors in accordance with the generated control commands to satisfy the one or more desired commands of the aircraft. 41. A flight control computer, One or more memory devices that store processor-executable instructions, A flight control computer comprising: one or more processors configured to execute the aforementioned instructions and cause the flight control computer to perform the methods performed by the computer described in any one of the clauses 28 to 40. 42. A non-temporary computer-readable medium storing computer-readable instructions, wherein when the computer-readable instructions are executed by at least one processor, the at least one processor causes the at least one processor to execute a method performed on a computer as described in any one of the clauses 28 to 40. 43. An aircraft having at least one flight control computer configured to carry out the computer-operated methods described in any one of the clauses 28 to 40. 44. A flight control system, At least one memory location that stores instructions, A flight control system comprising: at least one processor configured to execute the aforementioned instructions and carry out the methods performed on a computer as described in any one of the clauses 28 to 40. 45. An aircraft flight control system, At least one memory location that stores instructions, The system comprises at least one processor configured to execute the aforementioned instructions and perform one or more operations, wherein the operations are To determine one or more desired directives for an aircraft, Receiving engine information relating to at least one of the multiple electric propulsion units (EPUs) of the aircraft, wherein the engine information includes at least one temperature associated with the at least one EPU. Based on the received engine information, control commands are generated for multiple effectors of the aircraft. A flight control system comprising controlling the plurality of effectors according to the generated control commands to satisfy the one or more desired commands of the aircraft. 46. ​​The flight control system according to Clause 45, wherein the engine information further includes a time related to at least one temperature generated by the at least one EPU. 47. The flight control system according to Clause 46, wherein the time is generated by determining the current temperature of each component of a plurality of engine components for at least one EPU. 48. The flight control system according to Clause 47, wherein the time is generated by normalizing the determined temperature of one or more engine components having different temperature limits among the plurality of engine components using a time-based metric for temperature. 49. The flight control system according to Clause 48, wherein the aforementioned time is generated based on a prediction of the remaining time at the current torque rating associated with the engine of the EPU. 50. The flight control system according to any one of the clauses 45 to 49, wherein the engine information further includes the status of at least one engine associated with the plurality of EPUs. 51. The flight control system described in Clause 50, wherein the status includes one of active or inactive. 52. A flight control system according to any one of the clauses 45 to 51, wherein the generation of the control commands is further based on the maximum torque that the engine can generate under normal operating conditions. 53. A flight control system according to any one of clauses 45 to 52, wherein the generation of the control commands is further based on one or more aircraft conditions. 54. The flight control system described in Clause 53, wherein the one or more aircraft conditions include one or more of vehicle dynamics, flight conditions, or the status of at least one aircraft component. 55. The flight control system according to Clause 46, wherein the time includes the remaining time determined by the machine learning model of the engine. 56. The flight control system according to any one of the clauses 45 to 55, wherein the engine information further includes the temperature associated with the engine of the at least one EPU. 57. A non-temporary computer-readable medium storing one or more instructions, wherein when one or more instructions are executed by at least one processor, the medium causes the at least one processor to perform an operation, and the operation is To determine one or more desired directives for an aircraft, Receiving engine information relating to at least one of the multiple electric propulsion units (EPUs) of the aircraft, wherein the engine information includes at least one temperature associated with the at least one EPU. Based on the received engine information, control commands are generated for multiple effectors of the aircraft. A non-temporary computer-readable medium comprising controlling the plurality of effectors according to the generated control commands to satisfy the one or more desired commands of the aircraft. 58. A computer-readable medium storing computer-readable instructions, wherein when the computer-readable instructions are executed by at least one processor, the computer-readable medium causes at least one processor to perform the method described in any one of the clauses 28 to 40. 59. An aircraft engine, At least one memory location that stores instructions, The system comprises at least one processor configured to execute the aforementioned instructions and perform one or more operations, wherein the operations are Determining one or more temperatures associated with one or more components of the engine, An engine, comprising estimating the remaining time for the engine at the current power setting based on the one or more temperatures determined, wherein the estimated remaining time corresponds to a prediction of when the engine will reach one or more predetermined limits. 60. A method that is performed on a computer, To determine one or more desired directives for electric aircraft, Receiving engine information relating to at least one engine of at least one electric propulsion unit (EPU), To receive battery information for at least one battery pack, Based on the received engine information and battery information, control commands are generated for multiple effectors of the electric aircraft. A method performed by a computer, comprising controlling the plurality of effectors according to the generated control commands to satisfy the one or more desired commands of the electric aircraft.

[0191] The above description is provided for illustrative purposes only. It is not exhaustive and does not limit the invention to the exact forms or embodiments disclosed herein. Modifications and adaptations of the invention will be apparent to those skilled in the art from the examination herein and the practice of the disclosed embodiments disclosed herein.

[0192] The flowcharts and block diagrams in the figures illustrate examples of the architecture, functionality, and operation of possible implementations of systems, methods, and computer program products in various embodiments. In this regard, each block in a flowchart or block diagram may represent a module, segment, or portion of code containing one or more executable instructions for implementing a specified logical function(s). Note that in some alternative implementations, the functions described in a block may occur out of the order shown in the diagram. For example, two blocks shown consecutively may actually be executed substantially simultaneously, or blocks may sometimes be executed in reverse order depending on the functionality involved. Also note that each block in a block diagram or flowchart, and combinations of blocks in a block diagram or flowchart, may be implemented by a special-purpose hardware-based system that performs a specified function or operation, or a combination of special-purpose hardware and computer instructions.

[0193] Exemplary embodiments are described above with reference to flowcharts or block diagrams of methods, apparatus (systems), and computer program products. It will be understood that each block in a flowchart or block diagram, and combinations of blocks in a flowchart or block diagram, can be implemented by a computer program product or instructions on a computer program product. These computer program instructions may be provided to a processor of a computer or other programmable data processing device to generate a machine, thereby creating means for instructions executed via the processor of the computer or other programmable data processing device to implement the function / operation specified in the block(s) of the flowchart or block diagram.

[0194] These computer program instructions may also be stored in a computer-readable medium that can instruct one or more hardware processors of a computer, other programmable data processing device, or other device to function in a particular way, thereby forming a product containing instructions that implement functions / operations specified in a block(s) of a flowchart or block diagram.

[0195] Computer program instructions can also be loaded onto a computer, another programmable device, or another device so that a series of operational steps are performed (e.g., executed) on the computer, another programmable device, or another device, thereby providing a process for implementing a function / operation specified in a block(s) of a flowchart or block diagram.

[0196] Any combination of one or more computer-readable media may be used. Computer-readable media may be non-temporary computer-readable storage media. In the context of this document, computer-readable storage media may be any tangible media that contains or can store programs for use by, or in connection with, an instruction execution system, apparatus, or device.

[0197] Program code, as embodied on a computer-readable medium, may be transmitted using any suitable medium, including but not limited to wireless, wired, fiber optic cables, RF, IR, or any preferred combination thereof.

[0198] The features and advantages of this disclosure are evident from the detailed specification, and therefore the attached claims are intended to cover all systems and methods included in the true intent and scope of this disclosure. As used herein, the indefinite articles "a" and "an" mean "one or more." Similarly, the use of plural terms does not necessarily indicate plural unless it is clear in the given context. Words such as "and" or "or" mean "and / or" unless otherwise specifically indicated. As used herein, unless otherwise stated, "based on" may include depending on, being interdependent on, being associated with, being defined at least partially by, being influenced by, or responding to. As used herein, "related to" may include, being represented by, being indicated by, or being based on. Furthermore, since numerous modifications and variations are readily possible by examining this disclosure, it is not desirable to limit this disclosure to the exact configurations and operations illustrated and described, and therefore all suitable modifications and equivalents included in the scope of this disclosure may be applicable.

[0199] Other embodiments will be apparent to those skilled in the art from the examination of this specification and the practice of the implementations disclosed herein. The architectures and circuit layouts shown in the figures are intended for illustrative purposes only and are not intended to be limited to the specific configurations and circuit layouts described and illustrated herein. Furthermore, this specification and the examples are intended to be considered merely illustrative, and the true scope and spirit of the invention are given by the following claims. The above description is presented for illustrative purposes only. It is not exhaustive and does not limit the invention to the exact forms or embodiments disclosed herein. Modifications and adaptations of the invention will be apparent to those skilled in the art from the examination of this specification and the practice of the disclosed embodiments of the invention disclosed herein.

Claims

1. A method that is performed on a computer, To determine one or more desired directives for an aircraft, Based on the one or more desired directives and one or more aircraft conditions, determine at least one reference directive, Monitoring the energy state of multiple battery packs of the aircraft, wherein at least one first battery pack among the multiple battery packs is electrically isolated from at least one second battery pack among the multiple battery packs. Adjusting the at least one reference command based on the monitored energy state of the plurality of battery packs, Based on the adjusted reference command, generate control commands for multiple effectors of the aircraft, A method performed by a computer, comprising controlling the plurality of effectors according to the generated control commands in order to satisfy the one or more desired commands of the aircraft.

2. The method performed on a computer according to claim 1, wherein the monitored energy state includes the available energy of the plurality of battery packs.

3. The method performed by a computer according to claim 1 or 2, wherein the monitored energy state includes the remaining discharge time associated with one or more high-voltage channels.

4. A method performed on a computer according to any one of claims 1 to 3, wherein the monitored energy state includes the difference in energy states between at least two of the plurality of battery packs.

5. A method performed by a computer according to any one of claims 1 to 4, wherein the monitored energy state includes the difference in power consumption between at least a first engine and a second engine of a plurality of electric propulsion units of the aircraft.

6. The method performed by a computer according to claim 5, wherein the first engine has lower power consumption than the second engine, and the generated control command causes a reduction in power consumption from the second engine.

7. A method, performed by a computer according to any one of claims 1 to 6, wherein adjusting the at least one reference command is further based on the state of one or more engines of the aircraft's multiple electric propulsion units.

8. A method performed in a computer according to any one of claims 1 to 7, wherein adjusting the at least one reference command includes updating one or more reference commands to optimize the remaining discharge time across one or more high-voltage channels.

9. The method performed by a computer according to any one of claims 1 to 8, wherein the at least one reference command includes one or more of a reference engine command, a reference power command, a reference torque command, or a reference speed command.

10. The method performed by a computer according to any one of claims 1 to 9, wherein the one or more aircraft conditions include one or more of vehicle dynamics, flight conditions, or the status of at least one aircraft component.

11. A method performed on a computer controlling an aircraft, comprising a plurality of effectors including a plurality of electric propulsion units and a plurality of battery packs that supply power to the plurality of electric propulsion units, To determine one or more desired commands for the aircraft, Based on the aforementioned desired directive and one or more aircraft conditions, determine at least one reference directive, Monitoring the energy state of the plurality of battery packs, wherein at least one first battery pack among the plurality of battery packs is electrically isolated from at least one second battery pack among the plurality of battery packs. Adjusting the at least one reference command based on the monitored energy state of the plurality of battery packs, To generate control commands for the plurality of effectors based on the adjusted reference command, A method performed by a computer, comprising controlling the plurality of effectors according to the generated control commands in order to satisfy the one or more desired commands of the aircraft.

12. It is a flight control computer, One or more memory devices that store processor-executable instructions, A flight control computer comprising: one or more processors configured to execute the aforementioned instructions and cause the flight control computer to perform the method performed by the computer according to any one of claims 1 to 11.

13. A non-temporary computer-readable medium storing computer-readable instructions, wherein when the computer-readable instructions are executed by at least one processor, the at least one processor is instructed to execute the computer-executed method described in any one of claims 1 to 11.

14. An aircraft comprising at least one flight control computer configured to perform a computer-based method as described in any one of claims 1 to 11.

15. A flight control system, At least one memory that stores instructions, A flight control system comprising: at least one processor configured to execute the aforementioned instructions and carry out the computer-based method described in any one of claims 1 to 11.

16. An aircraft flight control system, At least one memory that stores instructions, The system comprises at least one processor configured to execute the aforementioned instructions and perform one or more operations, wherein the operations are To determine one or more desired directives for an aircraft, Based on the one or more desired directives and one or more aircraft conditions, determine at least one reference directive, Monitoring the energy state of multiple battery packs of the aircraft, wherein at least one first battery pack among the multiple battery packs is electrically isolated from at least one second battery pack among the multiple battery packs. Adjusting the at least one reference command based on the monitored energy state of the plurality of battery packs, Based on the adjusted reference command, generate control commands for multiple effectors of the aircraft, A flight control system comprising controlling the plurality of effectors according to the generated control commands to satisfy the one or more desired commands of the aircraft.

17. The flight control system according to claim 16, wherein the monitored energy state includes the available energy of the plurality of battery packs.

18. The flight control system according to claim 16 or 17, wherein the monitored energy state includes the remaining discharge time associated with one or more high-voltage channels.

19. The flight control system according to any one of claims 16 to 18, wherein the monitored energy state includes the difference in energy states between at least two of the plurality of battery packs.

20. The flight control system according to any one of claims 16 to 19, wherein the monitored energy state includes the difference in power consumption between at least a first engine and a second engine of a plurality of electric propulsion units of the aircraft.

21. The flight control system according to claim 20, wherein the first engine has lower power consumption than the second engine, and the generated control commands cause a reduction in power consumption from the second engine.

22. The flight control system according to any one of claims 16 to 21, wherein adjusting the at least one reference command is further based on the state of one or more engines of the aircraft's multiple electric propulsion units.

23. The flight control system according to any one of claims 16 to 22, wherein adjusting the at least one reference command includes updating one or more reference commands to optimize the remaining discharge time across one or more high-voltage channels.

24. The flight control system according to any one of claims 16 to 23, wherein the at least one reference command includes one or more of a reference engine command, a reference power command, a reference torque command, or a reference speed command.

25. The flight control system according to any one of claims 16 to 24, wherein the one or more aircraft conditions include one or more of vehicle dynamics, flight conditions, or the status of at least one aircraft component.

26. A non-temporary computer-readable medium storing one or more instructions, wherein when the one or more instructions are executed by at least one processor, the medium causes the at least one processor to perform an operation, and the operation is To determine one or more desired directives for an aircraft, Based on the one or more desired directives and one or more aircraft conditions, determine at least one reference directive, Monitoring the energy state of multiple battery packs of the aircraft, wherein at least one first battery pack among the multiple battery packs is electrically isolated from at least one second battery pack among the multiple battery packs. Adjusting the at least one reference command based on the monitored energy state of the plurality of battery packs, Based on the adjusted reference command, generate control commands for multiple effectors of the aircraft, A non-temporary computer-readable medium comprising controlling the plurality of effectors according to the generated control commands to satisfy the one or more desired commands of the aircraft.

27. A computer-readable medium storing computer-readable instructions, wherein when the computer-readable instructions are executed by at least one processor, the computer-readable medium causes the at least one processor to execute the method according to any one of claims 1 to 11.

28. A method that is performed on a computer, To determine one or more desired directives for an aircraft, Receiving engine information relating to at least one of the multiple electric propulsion units (EPUs) of the aircraft, wherein the engine information includes at least one temperature associated with the at least one EPU. Based on the received engine information, control commands are generated for multiple effectors of the aircraft. A method performed by a computer, comprising controlling the plurality of effectors according to the generated control commands in order to satisfy the one or more desired commands of the aircraft.

29. The method performed by a computer according to claim 28, wherein the engine information further includes a time related to at least one temperature generated by the at least one EPU.

30. The computer-based method according to claim 29, wherein the time is generated by determining the current temperature of each component of a plurality of engine components for at least one EPU.

31. The method performed by a computer according to claim 30, wherein the time is generated by normalizing the determined temperatures of the plurality of engine components, which are one or more engine components having different temperature limits, using a time-based metric for temperature.

32. The method performed by a computer according to claim 30, wherein the time is generated based on a prediction of the remaining time at the current torque rating associated with the engine of the EPU.

33. The method performed by a computer according to any one of claims 28 to 32, wherein the engine information further includes the status of at least one engine associated with the plurality of EPUs.

34. The method performed on a computer according to claim 33, wherein the status includes one of active or inactive.

35. A method, performed by a computer according to any one of claims 28 to 34, wherein the generation of the control command is further based on the maximum torque that the engine can generate under normal operating conditions.

36. A method by which the generation of the control command is performed by a computer according to any one of claims 28 to 35, further based on one or more aircraft conditions.

37. The method performed by a computer according to claim 36, wherein the one or more aircraft conditions include one or more of vehicle dynamics, flight conditions, or the status of at least one aircraft component.

38. The method performed on a computer according to claim 29, wherein the time includes the remaining time determined by the machine learning model of the engine.

39. A method performed by a computer according to any one of claims 28 to 38, wherein the engine information further includes the temperature associated with the engine of the at least one EPU.

40. A method for controlling an aircraft, comprising multiple effectors including multiple electric propulsion units (EPUs) and multiple battery packs that supply power to the multiple electric propulsion units, To determine one or more desired commands for the aircraft, Receiving engine information relating to at least one EPU among the plurality of EPUs, wherein the engine information includes at least one temperature associated with the at least one EPU. Based on the received engine information, control commands for the plurality of effectors are generated, A method comprising controlling the plurality of effectors in accordance with the generated control commands to satisfy the one or more desired commands of the aircraft.

41. It is a flight control computer, One or more memory devices that store processor-executable instructions, A flight control computer comprising one or more processors configured to execute the aforementioned instructions and cause the flight control computer to perform the method performed by the computer according to any one of claims 28 to 40.

42. A non-temporary computer-readable medium storing computer-readable instructions, wherein when the computer-readable instructions are executed by at least one processor, the at least one processor is instructed to execute the computer-executed method described in any one of claims 28 to 40.

43. An aircraft comprising at least one flight control computer configured to perform a computer-based method as described in any one of claims 28 to 40.

44. A flight control system, At least one memory that stores instructions, A flight control system comprising: at least one processor configured to execute the aforementioned instructions and carry out the computer-based method described in any one of claims 28 to 40.

45. An aircraft flight control system, At least one memory that stores instructions, The system comprises at least one processor configured to execute the aforementioned instructions and perform one or more operations, wherein the operations are To determine one or more desired directives for an aircraft, Receiving engine information relating to at least one of the multiple electric propulsion units (EPUs) of the aircraft, wherein the engine information includes at least one temperature associated with the at least one EPU. Based on the received engine information, control commands are generated for multiple effectors of the aircraft. A flight control system comprising controlling the plurality of effectors according to the generated control commands to satisfy the one or more desired commands of the aircraft.

46. The flight control system according to claim 45, wherein the engine information further includes a time related to at least one temperature generated by the at least one EPU.

47. The flight control system according to claim 46, wherein the aforementioned time is generated by determining the current temperature of each component of a plurality of engine components for at least one EPU.

48. The flight control system according to claim 47, wherein the time is generated by normalizing the determined temperatures of the plurality of engine components, which are one or more engine components having different temperature limits, using a time-based metric for temperature.

49. The flight control system according to claim 48, wherein the aforementioned time is generated based on a prediction of the remaining time at the current torque rating associated with the engine of the EPU.

50. The flight control system according to any one of claims 45 to 49, wherein the engine information further includes the status of at least one engine associated with the plurality of EPUs.

51. The flight control system according to claim 50, wherein the status includes one of active or inactive.

52. The flight control system according to any one of claims 45 to 51, wherein the generation of the control commands is further based on the maximum torque that the engine can generate under normal operating conditions.

53. The flight control system according to any one of claims 45 to 52, wherein the generation of the control commands is further based on one or more aircraft conditions.

54. The flight control system according to claim 53, wherein the one or more aircraft conditions include one or more of vehicle dynamics, flight conditions, or the status of at least one aircraft component.

55. The flight control system according to claim 46, wherein the time includes the remaining time determined by the machine learning model of the engine.

56. The flight control system according to any one of claims 45 to 55, wherein the engine information further includes the temperature associated with the engine of the at least one EPU.

57. A non-temporary computer-readable medium storing one or more instructions, wherein when the one or more instructions are executed by at least one processor, the medium causes the at least one processor to perform an operation, and the operation is To determine one or more desired directives for an aircraft, Receiving engine information relating to at least one of the multiple electric propulsion units (EPUs) of the aircraft, wherein the engine information includes at least one temperature associated with the at least one EPU. Based on the received engine information, control commands are generated for multiple effectors of the aircraft. A non-temporary computer-readable medium comprising controlling the plurality of effectors according to the generated control commands to satisfy the one or more desired commands of the aircraft.

58. A computer-readable medium storing computer-readable instructions, wherein when the computer-readable instructions are executed by at least one processor, the at least one processor is instructed to perform the method according to any one of claims 28 to 40.

59. It is an aircraft engine, At least one memory that stores instructions, The system comprises at least one processor configured to execute the aforementioned instructions and perform one or more operations, wherein the operations are Determining one or more temperatures associated with one or more components of the engine, An engine comprising estimating the remaining time for the engine at the current power setting based on the one or more temperatures determined, wherein the estimated remaining time corresponds to a prediction of when the engine will reach one or more predetermined limits.

60. A method that is performed on a computer, To determine one or more desired directives for electric aircraft, Receiving engine information relating to at least one engine of at least one electric propulsion unit (EPU), To receive battery information for at least one battery pack, Based on the received engine information and battery information, control commands are generated for multiple effectors of the electric aircraft. A method performed by a computer, comprising controlling the plurality of effectors according to the generated control commands to satisfy the one or more desired commands of the electric aircraft.

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