System and method for flight control of aircraft

Abstract

Aspects of the present disclosure generally relate to systems and methods for flight control of aircraft driven by an electric propulsion system, as well as other types of aircraft. In some embodiments, a flight control system of an aircraft is disclosed, the flight control system configured to: determine one or more desired commands for an electric aircraft; determining at least one reference command based on the desired command and one or more aircraft conditions; monitoring an energy state of a plurality of battery packs, wherein at least a first battery pack of the plurality of battery packs is electrically isolated from at least a second battery pack of the plurality of battery packs; adjusting the at least one reference command based on the monitored energy states of the plurality of battery packs; generating a control command for the plurality of actuators based on the adjusted at least one reference command; and controlling the plurality of actuators to satisfy the one or more desired commands of the electric aircraft according to the generated control commands.

Description

Cross-references to related applications

[0001] This disclosure claims priority to U.S. Provisional Application No. 63 / 512,784 (Agent No. 16499.6006-00000), filed July 10, 2023, entitled “SYSTEMS AND METHOD FOR FLIGHT CONTROL OF EVTOL AIRCRAFT,” the contents of which are incorporated herein by reference in their entirety for all purposes. Technical Field

[0002] This disclosure generally relates to powered aircraft. More specifically, but not limited to, this disclosure relates to innovations in aircraft powered by electric propulsion systems. Certain aspects of this disclosure generally relate to systems and methods for flight control of aircraft powered by electric propulsion systems and other types of aircraft, as well as for flight control of aircraft in flight simulators and video games. Other aspects of this disclosure generally relate to improvements in flight control systems and methods that provide specific advantages in powered aircraft and can be used in other types of aircraft. Background Technology

[0003] The inventors have recognized several problems that may be associated with flight control of aircraft, including tiltrotor aircraft using electric or hybrid electric propulsion systems (hereinafter referred to as electric propulsion units or "EPUs"). Vertical takeoff and landing (VTOL) aircraft are those capable of taking off, landing, and hovering vertically, thus enabling direct transport of passengers to their destinations. Helicopters are VTOL aircraft that generate lift entirely through their rotors. Some VTOL aircraft have wings and a propulsion system that allows the wings to provide the lift required during forward flight. Some winged VTOL aircraft use a separate propulsion system that uses vertical thrust during takeoff and landing and forward thrust during cruise. Other winged 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 both vertical and forward flight. Many electric VTOL aircraft include movable electric propulsion units, where the thrust vector of the propulsion unit can be changed, for example, from an upward direction for vertical lift to a forward direction for forward flight. Many electric VTOL aircraft are overacted because the actuators have more degrees of freedom than the translational degrees of freedom. Control allocation is the problem of distributing control forces among multiple actuators in an overacted system to optimize energy and manage the temperature of different aircraft components. Electric VTOL aircraft typically contain more propulsion units and other actuators compared to conventional aircraft, and these propulsion units and other actuators significantly influence multiple control axes. Therefore, electric VTOL aircraft may present greater control allocation challenges compared to conventional aircraft. Summary of the Invention

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

[0005] One aspect of this disclosure includes a method comprising: determining one or more desired commands for an aircraft; determining at least one reference command based on the one or more desired commands and one or more aircraft conditions; monitoring the energy state of a plurality of battery packs of an electric aircraft, wherein at least a first battery pack of the plurality of battery packs is electrically isolated from at least a second battery pack of the plurality of battery packs; adjusting the at least one reference command based on the monitored energy state of the plurality of battery packs; generating control commands for a plurality of actuators of the electric aircraft based on the adjusted at least one reference command; and controlling the plurality of actuators according to the generated control commands to satisfy the one or more desired commands of the electric aircraft.

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

[0007] Another aspect of this disclosure includes a non-transitory computer-readable medium storing one or more instructions that, when executed by at least one processor, cause the at least one processor to perform operations including: determining one or more desired commands for an electric aircraft; determining at least one reference command based on the one or more desired commands and one or more aircraft conditions; monitoring the energy state of a plurality of battery packs of the electric aircraft, wherein at least a first battery pack of the plurality of battery packs is electrically isolated from at least a second battery pack of the plurality of battery packs; adjusting the at least one reference command based on the monitored energy state of the plurality of battery packs; generating control commands for a plurality of actuators of the electric aircraft based on the adjusted at least one reference command; and controlling the plurality of actuators according to the generated control commands to satisfy the one or more desired commands of the electric aircraft.

[0008] Another aspect of this disclosure includes a method comprising: determining one or more desired commands for an electric aircraft; receiving engine information for 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 commands for a plurality of actuators of the electric aircraft based on the received engine information; and controlling the plurality of actuators according to the generated control commands to satisfy the one or more desired commands of the electric aircraft.

[0009] Another aspect of this disclosure includes a system comprising at least one processor configured to: determine one or more desired commands for an electric aircraft; receive engine information for 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; generate control commands for a plurality of actuators of the electric aircraft based on the received engine information; and control the plurality of actuators according to the generated control commands to satisfy the one or more desired commands of the electric aircraft.

[0010] Another aspect of this disclosure includes a non-transitory computer-readable medium storing one or more instructions that, when executed by at least one processor, cause the at least one processor to perform operations including: determining one or more desired commands for an electric aircraft; receiving engine information for 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 commands for a plurality of actuators of the electric aircraft based on the received engine information; and controlling the plurality of actuators according to the generated control commands to satisfy the one or more desired commands of the electric aircraft.

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

[0012] Another aspect of this disclosure includes a method comprising: determining one or more desired commands for an electric aircraft; receiving engine information for at least one EPU; receiving battery information for at least one battery pack; generating control commands for a plurality of actuators of the electric aircraft based on the received engine information and battery information; and controlling the plurality of actuators according to the generated control commands to satisfy the one or more desired commands of the electric aircraft. Attached Figure Description

[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 plan 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 shown.

[0017] Figure 5 An exemplary power connection in a VTOL aircraft consistent with the disclosed embodiments is shown.

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

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

[0020] Figure 8 This is a schematic diagram of a flight control signaling architecture for controlling control surfaces and associated actuators, consistent with the disclosed embodiments.

[0021] Figure 9A , Figure 9B , Figure 9C , Figure 9D and Figure 9E An exemplary top plan view of a VTOL aircraft consistent with the disclosed embodiments is shown.

[0022] Figure 10 A functional block diagram of an exemplary control system for an electric VTOL aircraft consistent with the disclosed embodiments is shown.

[0023] Figure 11 A functional block diagram of an exemplary energy optimization method according to some embodiments is shown.

[0024] Figure 12 Exemplary methods for energy optimization are shown according to some embodiments.

[0025] Figure 13A , Figure 13B , Figure 13C and Figure 13D Exemplary battery pack failure scenarios are shown according to some embodiments.

[0026] Figure 14 , Figure 15 , Figure 16 , Figure 17 and Figure 18 Exemplary scenarios for energy optimization are shown according to some embodiments.

[0027] Figure 19An exemplary temperature versus time graph is shown according to some embodiments.

[0028] Figure 20 An exemplary meter related to engine temperature is shown according to some embodiments.

[0029] Figure 21 Exemplary temperature thresholds related to engine temperature are shown according to some embodiments.

[0030] Figure 22 A block diagram of an exemplary control system for an aircraft according to some embodiments is shown.

[0031] Figure 23 An exemplary method for engine temperature management is shown according to some embodiments.

[0032] Figure 24 A block diagram of an exemplary control system including energy optimization and engine temperature optimization functions according to some embodiments is shown.

[0033] Figure 25A and Figure 25B Exemplary scenarios for implementing thermal management and energy optimization according to some embodiments are shown. Detailed Implementation

[0034] This disclosure discloses systems, components, and techniques primarily for use with aircraft. The aircraft can be a manned aircraft, an unmanned aircraft (e.g., a UAV), a drone, a helicopter, and / or an airplane. An aircraft comprises a fuselage and one or more components configured to enable flight (e.g., wings, tail, propeller, actuator, engine or motor, propulsion unit, airframe, actuator). The aircraft can include any configuration comprising at least one propeller. In some embodiments, the aircraft is driven (e.g., provides thrust) by one or more electric propulsion systems (hereinafter referred to as "electric propulsion unit" or "EPU"), which may include at least one engine, at least one rotor, at least one propeller, or any combination thereof. The aircraft can be fully electric, hybrid, or fuel-powered. For example, in some embodiments, the aircraft is a tiltrotor aircraft configured for frequent (e.g., more than 50 flights per workday), short-duration (e.g., less than 100 miles per flight) flights over, into, and out of densely populated areas. Aircraft can be configured to carry 4 to 6 passengers or commuters who desire a comfortable experience with low noise and low vibration. Therefore, it is desirable to control aircraft components in a way that optimizes aircraft resources to improve aircraft performance (e.g., increasing safety, energy efficiency, passenger comfort, load capacity, or structural integrity).

[0035] The disclosed embodiments provide new and improved configurations for some aircraft components not observed in conventional aircraft, and / or design standards for identifying components that differ from those in conventional aircraft. This combination of alternative configurations and design standards addresses the shortcomings and challenges of conventional components, resulting in the various configurations and designs disclosed herein for components of propulsion-driven aircraft (e.g., electric or hybrid electric aircraft).

[0036] In some embodiments, the propulsion-driven aircraft of this disclosure can be designed to perform both vertical takeoff and landing, as well as conventional takeoff and landing, wherein the distributed propulsion system enables vertical flight, horizontal and lateral flight, and transitions (e.g., transitions between vertical and horizontal flight). The aircraft can generate thrust by supplying high-voltage (HV) electricity to multiple engines of the distributed propulsion system, which may include components that convert the high-voltage electricity into mechanical shaft power to rotate the propeller.

[0037] Embodiments may include an electric engine (e.g., a motor) connected to an onboard power source, which may include devices capable of storing energy, such as batteries or capacitors, and may optionally include one or more systems for utilizing or generating electricity, such as fuel-powered generators or solar panel arrays. In some embodiments, the aircraft may include a hybrid-electric aircraft that uses at least one of an electric or fuel-based energy source to power a distributed propulsion system. In some embodiments, the aircraft may be powered by one or more batteries, an internal combustion engine (ICE), a generator, a turbine engine, or a ducted fan.

[0038] The engines can be mounted directly to the wing or to one or more booms attached to the wing. The amount of thrust generated by each engine can be controlled by torque commands sent to each engine via a digital communication interface by the flight control system (FCS). Embodiments may include a front engine (and associated propeller) capable of changing its direction or tilting.

[0039] The engine can also rotate the propeller in either a clockwise or counterclockwise direction. In some embodiments, the difference in propeller rotation direction can be achieved using the direction of engine rotation. In other embodiments, the engine can rotate entirely in the same direction, and different propeller rotation directions can be achieved using gear mechanisms.

[0040] In some embodiments, an aircraft may have a number of engines in various combinations of nose and tail engine configurations. A nose engine can be considered as an engine primarily oriented toward the leading edge of the wing. A tail engine can be considered as an engine primarily oriented toward the trailing edge of the wing. For example, an aircraft may have six nose and six tail engines, five nose and five tail engines, four nose and four tail engines, three nose and three tail engines, two nose and two tail engines, or any other combination of nose and tail engines, including embodiments with unequal numbers of nose and tail engines.

[0041] In some embodiments, for vertical takeoff and landing (VTOL) missions, the front and rear engines can provide vertical thrust during takeoff and landing. During the forward flight phase, the front engine can provide horizontal thrust, while the propellers of the rear engines can retract to a fixed position to minimize drag. The rear engines can be actively retracted using position monitoring.

[0042] The transition from vertical to horizontal flight and from horizontal to vertical flight can be achieved via a tilting propeller subsystem. The tilting propeller subsystem can redirect thrust between the primary vertical direction during the vertical flight phase (e.g., hovering phase) and the horizontal or near-horizontal direction during the forward flight cruise phase, based on the tilt of one or more propellers (e.g., determining the directionality of one or more propellers). A variable pitch mechanism can change the blade collective angle of the propeller hub assembly of the front engine for operation during flight phases such as hovering, transition, and cruise. Vertical lift can be thrust in the primary vertical direction (e.g., during hovering). Horizontal thrust can be thrust in the primary horizontal direction (e.g., during cruise). In some embodiments, the phase of flight (e.g., hovering, cruise, forward flight, takeoff, landing, transition to forward flight, or transition 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., the pitch angle of the aircraft), roll angle (e.g., the roll angle of one or more propellers), roll angle, rotational speed (e.g., the rotational speed of the propellers), torque value, pilot command, sensor measurement (e.g., accelerometer measurement), or any other value indicating the current (e.g., experienced) or requested (e.g., commanded) state of at least a portion of the aircraft.

[0043] In some embodiments, during conventional takeoff and landing (CTOL) missions, the front engine can provide horizontal thrust for wing-borne takeoff, cruise, and landing, and the wing can provide vertical lift. In some embodiments, the tail engine may not be used to generate thrust during CTOL missions, and the tail propeller may be retracted in place. In other embodiments, the tail engine may be used at reduced power to shorten the length of CTOL takeoff or landing.

[0044] As described above, embodiments of an aircraft may include numerous movable structural flight elements that allow a pilot to safely control the aircraft. Flight control surfaces (e.g., flaps, ailerons, elevators, rudders, etc.) are critical for controlling the aircraft's positioning. Changing the orientation of these surfaces alters the airflow and pressure distribution around the aircraft, enabling the pilot to control the aircraft's movement along three axes of rotation. Similarly, propeller rotation and directional control can provide lift support (e.g., the lift required for vertical takeoff, landing, and hovering) and can provide the forward thrust required for the aircraft to fly in the air. The movement of each of these flight elements is critical to the aircraft's safety and stability.

[0045] Reference will now be made in detail to exemplary embodiments, examples of which are illustrated in the accompanying drawings. The following description refers to the accompanying drawings, in which, unless otherwise stated, the same numerals in different figures represent the same or similar elements. The embodiments set forth in the following description of the exemplary embodiments do not represent all embodiments consistent with this disclosure. Instead, they are merely examples of apparatuses and methods consistent with aspects relevant to the subject matter recited in the appended claims.

[0046] Figure 1 This is an illustration of a perspective view of an exemplary VTOL aircraft consistent with the disclosed embodiments. Figure 2 This is another illustration of a perspective view of an exemplary VTOL aircraft in an alternative configuration consistent with embodiments of this disclosure. Figure 1 and Figure 2 VTOL aircraft 100 and 200, consistent with embodiments of this disclosure, are shown in cruise configuration and vertical takeoff, landing, and hovering configurations (also referred to herein as "lift" configurations). Figure 1 and Figure 2Corresponding components may have similar reference numerals and refer to similar components of aircraft 100, 200. Aircraft 100, 200 may include fuselages 102, 202, wings 104, 204 mounted to fuselages 102, 202, and one or more rear stabilizers 106, 206 mounted to the rear of fuselages 102, 202. Multiple lift propellers 112, 212 may be mounted to wings 104, 204 and may be configured to provide lift for vertical takeoff, landing, and hovering. Multiple tilt propellers 114, 214 may be mounted to wings 104, 204 and may tilt (e.g., be configured to tilt or change direction) between a lift configuration and a cruise configuration, in which the multiple tilt propellers provide a portion of the lift required for vertical takeoff, landing, and hovering, such as... Figure 2 As shown, in the cruise configuration, the plurality of tilting propellers provide forward thrust to the aircraft 100 for horizontal flight, such as... Figure 1 As shown. As used herein, a tilt propeller lift configuration refers to any tilt propeller direction in which tilt propeller thrust primarily provides lift to the aircraft, while a tilt propeller cruise configuration refers to any tilt propeller direction in which tilt propeller thrust primarily provides forward thrust to the aircraft.

[0047] In some embodiments, lift propellers 112, 212 can be configured to provide lift only, with all horizontal propulsion provided by tilting propellers. For example, lift propellers 112, 212 can be configured in a fixed position and can generate thrust only during takeoff, landing, and hovering phases of flight. Meanwhile, tilting propellers 114, 214 can be tilted upwards into a lift configuration in which thrust from propellers 114, 214 is directed downwards to provide additional lift.

[0048] For forward flight, tilting propellers 114 and 214 can tilt from their lift configuration to their cruise configuration. In other words, the direction of tilting propellers 114 and 214 can change from a direction where tilting propeller thrust is directed downwards (to provide lift during vertical takeoff, landing, and hovering) to a direction where tilting propeller thrust is directed backwards (to provide forward thrust to aircraft 100 and 200). The tilting propeller assembly for a specific electric engine can tilt about an axis of rotation defined by the mounting point connecting the boom and the electric engine. When aircraft 100 and 200 are in fully forward flight, lift can be provided entirely by wings 104 and 204. Meanwhile, in cruise configuration, lift propellers 112 and 212 can be shut off. The blades 120 and 220 of lift propellers 112 and 212 can be maintained in a low-drag position for aircraft cruise. In some embodiments, the lift propellers 112 and 212 may each have two blades 120 and 220, respectively, which may be locked in a minimum drag position, for example, during aircraft cruise, in which one blade is directly in front of the other blade, such as... Figure 1 As shown. In some embodiments, the lift propellers 112, 212 have more than two blades. In some embodiments, the tilting propellers 114, 214 may contain more blades 116, 216 than the lift propellers 112, 212. For example, as Figure 1 and Figure 2 As shown, lift propellers 112 and 212 may each contain, for example, two blades, while tilt propellers 114 and 214 may each contain more blades, such as the five blades shown. In some embodiments, each of tilt propellers 114 and 214 may have two to five blades, and possibly more, depending on the aircraft's design considerations and requirements.

[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 across the entire aircraft). At least a portion of the lift propellers 112, 212 may be located at the rear of the wings 104, 204 (e.g., from a bird's-eye view, the propeller's rotation point is behind the wing), and at least a portion of the jib propellers 114, 214 may be located at the front of the wings 104, 204 (e.g., from a bird's-eye view, the propeller's rotation point is in front of the wing). In some embodiments, all of the lift propellers 112, 212 may be located at the rear of the wings 104, 204, and all of the jib propellers 114, 214 may be located at the front of the wings 104, 204. According to some embodiments, all the lift propellers 112, 212 and the jib propellers 114, 214 may be mounted to the wings—that is, no lift propeller or jib propeller may be mounted to the fuselage. In some embodiments, the lift propellers 112 and 212 may all be located at the rear of the wings 104 and 204, and the tilt propellers 114 and 214 may all be located at the front of the wings 104 and 204. According to some embodiments, all lift propellers 112 and 212 and tilt propellers 114 and 214 may be positioned inside the ends of the wings 104 and 204.

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

[0051] In some embodiments, aircraft 100, 200 may include, for example, a wing on each side of fuselage 102, 202 or a single wing extending across the aircraft. According to some embodiments, at least one wing 104, 204 is a high wing mounted to the upper side of fuselage 102, 202. According to some embodiments, the wing includes control surfaces, such as flaps, ailerons, and / or flaperons (e.g., configured to perform the functions of both flaps and ailerons). According to some embodiments, 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 combinations of rudder-elevator. The wing may have any suitable design for providing lift, directionality, stability, and / or any other characteristics beneficial to the aircraft. In some embodiments, the wing has a tapered leading edge.

[0053] In some embodiments, the lift propellers 112, 212 or the tilt propellers 114, 214 may be tilted relative to at least one other lift propeller 112, 212 or tilt propeller 114, 214, wherein tilting refers to the relative orientation of the rotation axis of the lift propeller / tilt propeller about a line parallel to the forward and backward direction, similar to the roll degree of freedom of an aircraft.

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

[0055] Figure 3 This is an illustration of a top plan view of an exemplary VTOL aircraft consistent with embodiments of this disclosure. The aircraft 300 shown in the figure may be located in… Figure 1 and Figure 2The diagram shows top-down plan views of aircraft 100 and 200. As discussed herein, aircraft 300 may include twelve electric propulsion systems distributed across aircraft 300. In some embodiments, the distribution of electric propulsion systems may include six forward electric propulsion systems 314 and six tail electric propulsion systems 312 mounted on booms at the forward and tail ends of the main wing 304 of aircraft 300. In some embodiments, the forward electric propulsion systems may be mounted to the wing 304 via booms 322. In some embodiments, the tail electric propulsion systems may be mounted to the wing 304 via booms 324. In some embodiments, the length of the boom 324 from the wing 304 to the rear end of the lifting propeller (part of the electric propulsion system 312) may include similar rear end lengths of boom 324 across multiple rear ends of the boom. In some embodiments, the rear end lengths of the booms may vary, for example, across six rear ends of the boom. Furthermore, Figure 3 An exemplary embodiment of a VTOL aircraft 300 is depicted, having a front propeller (part of an electric propulsion system 314) in a horizontal position for horizontal flight and a tail propeller blade 320 in a retracted position for forward flight.

[0056] Figure 4 This 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, respectively, in… Figure 1 , Figure 2 and Figure 3The diagram shows top-down plan views of aircraft 100, 200, and 300. Aircraft 400 may include six forward electric propulsion systems, three of which have a clockwise (CW) type 424, while the remaining three have a CCW type 426. In some embodiments, three tail electric propulsion systems may have a counter-clockwise (CCW) type 428, while the remaining three tail electric propulsion systems have a CW type 430. Some embodiments may include an aircraft 400 having four forward electric propulsion systems and four tail electric propulsion systems, each having two CW types and two CCW types. In some embodiments, aircraft 400 may include a fuselage 402, wings 404 mounted to the fuselage 402, and one or more rear stabilizers 406 mounted to the rear of the fuselage 402. In some embodiments, each forward electric propulsion system may include a propeller blade 416. In some embodiments, each tail-end electric propulsion system may include propeller blades 420. In some embodiments, the electric propulsion system may be mounted to the wing 404 via a boom 422. In some embodiments, the propellers may rotate in opposite directions relative to adjacent propellers to eliminate torque steer experienced by the aircraft fuselage or wing due to propeller rotation. In some embodiments, the difference in rotation direction may be achieved using the direction of engine rotation. In other embodiments, the engines may all rotate in the same direction, and different propeller rotation directions may be achieved using gear mechanisms.

[0057] Some embodiments may include an aircraft 400 having a front electric propulsion system and a rear electric propulsion system, wherein the amount of CW type 424 and CCW type 426 is not equal in the front electric propulsion system, in the rear electric propulsion system, or in both the front and rear electric propulsion systems.

[0058] Figure 5 An example eVTOL aircraft according to an embodiment of this disclosure is shown. 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) electricity. The HVPS system may be configured to supply power on the aircraft 500 and to distribute the stored electrical energy to other systems on the aircraft 500 (e.g., via high-voltage channels, via high-voltage buses), including an electric propulsion system (EPS) for converting the electricity into mechanical rotating shaft power to generate thrust. Figure 5As shown, the HVPS system of aircraft 500 may include six battery packs 520 (numbered B1-B6 from left to right) mounted in battery trays in the wings of 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 can power one or more electric engines 510 (numbered E1 to E12).

[0059] In some embodiments, a single battery pack 520 may be electrically connected to and power a plurality of electric motors 510. For example, in some embodiments, the battery pack 520 may power an electric motor 510 on either side of a longitudinal axis. In some embodiments, the battery pack 520 may power an electric motor 510 on either side of a horizontal axis (e.g., along one or more wings of an aircraft). In some embodiments, such as Figure 5 As shown, battery pack 520 can power two diagonally opposite 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, in the event of a loss in battery pack 520, the impact on roll or pitch moments can be reduced due to optimized lift loss. In some embodiments, battery pack 520 can power different arrangements of electric engines 510 to reduce roll, pitch, or yaw moments that may be caused by a loss in battery pack 520. For example, in some embodiments, battery pack 520 can be connected to electric engines 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 crosslinker 530 (e.g., a high-voltage bus) having a fuse capable of pairing two or more battery packs 520. Through this crosslinker, power for the electric motor 510 can be shared among the paired battery packs 520. Therefore, multiple battery packs 520 can simultaneously power multiple electric motors 510. This arrangement provides redundancy and avoids single points of failure, as each paired battery 520 can act as a backup for the others. In the event of a battery pack 520 failure, one or more connected battery packs 520 can continue to power the electric motor 510 connected to the failed battery pack.

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

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

[0063] In some embodiments, more than two 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 can power six electric engines 510. In some embodiments, four battery packs 520 may be cross-linked. Thus, in some embodiments, four battery packs 520 can power eight electric engines 510. In some embodiments, different arrangements of the battery packs 520 and cross-linking members can be selected to optimally optimize aircraft power requirements, system redundancy, and fault tolerance.

[0064] In some embodiments, each battery pack may contain an HV distribution unit and a battery management system housed within 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 prevent fault conditions (e.g., using fuses). As described further below, the BMS can communicate with various systems both inside and outside the HVJB. The BMS may include a battery management unit (BMU) that receives voltage, current, resistance, and temperature sensing signals from the battery stack assembly and / or the HV distribution unit.

[0065] The BMU can monitor the output current of each connected load. Even when the BMU is not installed in the aircraft 100, it can continuously monitor the battery status. By monitoring the battery pack, the BMU can prevent conditions that adversely affect safety or performance, such as overvoltage, undervoltage, overtemperature, undertemperature, electrical isolation loss, short circuit, overcurrent, etc. Additionally, 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 (SOE), and state of temperature (SOT). In some embodiments, the BMU can be configured to perform calculations of usable energy (e.g., available energy) based on one or more of SOC, SOH, SOP, SOE, or SOT. For example, consistent with the disclosed embodiments, the BMU can perform calculations of usable energy for each battery pack. The BMU also controls and monitors bus precharging, provides fuse and contactor commands, and communicates with various systems internal and external to the HVJB. Furthermore, the BMU can communicate with the aircraft's switches and flight control system and modify operations based on received commands.

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

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

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

[0069] Some embodiments may include an electric propulsion system 602 that includes an electric engine subsystem 604 that receives signals from and transmits signals to a flight control system 612. In some embodiments, the flight control system (FCS) 612 may include a flight control computer (FCC) capable of sending commands to and receiving status and data from the electric engine subsystem 604 using controller area network (“CAN”) data bus signals. Consistent with the disclosed embodiments, the FCC may include means configured to perform one or more operations (e.g., computational operations) on the aircraft, such as at least one processor and a memory component that may store instructions executable by at least one processor to perform operations. It should be understood that while CAN data bus signals are used between the flight control computer and the electric engine, some embodiments may include any form of communication capable of sending and receiving data from the flight control computer to the electric engine. Some embodiments may include an electric motor subsystem 604 capable of receiving operating parameters from and transmitting operating parameters to the FCC in the FCS 612, such operating parameters including speed, voltage, current, torque, temperature, vibration, propeller position, and / or any other values ​​of operating parameters.

[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 electric engine subsystem 604 of the tilt propeller. The tilt propeller system (TPS) 614 may include devices capable of transmitting operating parameters to the electric engine subsystem 604 and hinged to the orientation of the propeller subsystem 606 to redirect the thrust of the tilt propeller by changing the orientation of the propeller subsystem 606 using mechanical components (e.g., gearbox assembly, linear actuators, and any other configured components) during various phases of flight. 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 control surfaces and their associated actuators in an exemplary VTOL aircraft. Figure 7 This is an illustration of a top plan view of an exemplary VTOL aircraft consistent with embodiments of this disclosure. In addition to the references above... Figure 3 Apart from the aircraft components described, the aircraft 700 shown in the figure can be respectively located in... Figure 1 and Figure 2 The diagram shows top-view plans of aircraft 100 and 200. In aircraft 700, in addition to the propeller blades discussed previously, the control surface may also include flaperon 712 and rudder 714. The flaperon 712 may combine the functions of one or more flaps, one or more ailerons, and / or one or more spoilers. The rudder 714 may combine the functions of one or more rudders and / or one or more elevators. Alternatively or additionally, the control surface may include separate rudders and elevators. In aircraft 700, in addition to the electric propulsion system discussed previously, the actuators may also include control surface actuators (CSAs) associated with the flaperon 712 and rudder 714, as referenced below. Figure 8 Further discussion.

[0072] Figure 8 This is a schematic diagram of a flight control signaling architecture 800 for controlling control surfaces and associated actuators, consistent with the disclosed embodiments. Although Figure 7 The diagram shows twelve EPU inverters and associated propeller blades, six tilt propeller actuators (TPAC), six battery management systems (BMS), four flaps and associated control surface actuators (CSA), and six directional elevators and associated CSA; however, aircraft according to various embodiments may have any suitable number of these various elements. Figure 8As shown, the control surfaces and actuators can be controlled by a combination of four flight control computers (FCCs)—left FCC, channel A (L FCC-A) 801, left FCC, channel B (L FCC-B) 802, right FCC, channel A (RFCC-A) 803, and right FCC, channel B (RFCC-B) 804—but any other suitable number of FCCs can be used. Each FCC can control all control surfaces and actuators individually, or in any combination thereof. In some embodiments, each FCC can contain one or more hardware computing processors. In some embodiments, each FCC can 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 calculations required to control the control surfaces and actuators can be performed by a single flight control computer on a single computing thread.

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

[0074] Figures 9A to 9EThis is an illustration of a top plan view of an exemplary VTOL aircraft consistent with embodiments of this disclosure. Numerous design considerations (cost, weight, size, performance capabilities, etc.) may exist that could influence the number and / or combination of tiltrotor and lift propellers in a VTOL aircraft. As further described below, the number and orientation of aircraft components (e.g., actuators or actuators) can affect how energy is used and the temperature of the various aircraft components. Therefore, the flight control system can adjust the aircraft components in certain ways (e.g., those discussed in the disclosed embodiments) to control the aircraft in a manner that optimizes energy use and maintains the various aircraft components at optimal temperatures.

[0075] Figure 9A An arrangement of an electric propulsion system 900 consistent with an embodiment of this disclosure is shown. Reference Figure 9A The aircraft shown in the figure may be a top plan view of an exemplary aircraft (e.g., a VTOL aircraft). The aircraft may contain twelve 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 tail electric propulsion systems 907, 908, 909, 910, 911, 912. In some embodiments, the six forward electric propulsion systems may be operatively connected to a tilting propeller, and the six tail electric propulsion systems may be operatively connected to a lift propeller. In other embodiments, the six forward electric propulsion systems and multiple tail electric propulsion systems may be operatively connected to a tilting propeller, and the remaining tail electric propulsion systems may be operatively connected to a lift propeller. In other embodiments, all forward and tail electric propulsion systems may be operatively coupled to a tilting propeller.

[0076] Figure 9B An alternative arrangement of the electric propulsion system 940 consistent with embodiments of this disclosure is shown. Reference Figure 9BThe aircraft shown in the figure may be a top plan 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 tail electric propulsion systems 917, 918, 919, 920. In some embodiments, the four forward electric propulsion systems may be operatively connected to a tilting propeller, and the four tail electric propulsion systems may be operatively connected to a lift propeller. In other embodiments, the four forward electric propulsion systems and multiple tail electric propulsion systems may be operatively connected to a tilting propeller, and the remaining tail electric propulsion systems may be operatively connected to a lift propeller. In other embodiments, all forward and tail electric propulsion systems may be operatively coupled to a tilting propeller.

[0077] Figure 9C An alternative arrangement of the electric propulsion system 950 consistent with embodiments of this disclosure is shown. Reference Figure 9C The aircraft shown in the figure may be a top plan 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 group of four electric propulsion systems 921, 922, 923, 924 coplanar in a first plane and a second group of two electric propulsion systems 925, 926 coplanar in a second plane. In some embodiments, the first group of electric propulsion systems 921-924 may be operatively connected to a tilting propeller, and the second group of electric propulsion systems 925, 926 may be operatively connected to a lift propeller. In other embodiments, the first group of electric propulsion systems 921-924 and the second group of tail-mounted electric propulsion systems 925, 926 may all be operatively connected to the tilting propeller.

[0078] Figure 9D An alternative arrangement of the electric propulsion system 960 consistent with embodiments of this disclosure is shown. Reference Figure 9D The aircraft shown in the figure may be a top plan view of an exemplary aircraft (e.g., a VTOL aircraft). The aircraft may include four electric propulsion systems distributed across 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 electric propulsion systems may be operatively connected to tilting propellers.

[0079] Figure 9E An alternative arrangement of the electric propulsion system 970 consistent with embodiments of this disclosure is shown. Reference Figure 9EThe aircraft shown in the figure may be a top plan view of an exemplary aircraft (e.g., a VTOL aircraft). The aircraft may include six electric propulsion systems distributed across the aircraft. For example, in some embodiments, the aircraft may include four forward electric propulsion systems 931, 932, 933, 934 operably connected to a tilting propeller and two tail electric propulsion systems 935 and 936 operably connected to a lift propeller. In some embodiments, the aircraft may include ten electric propulsion systems distributed across the aircraft. For example, in some embodiments, the aircraft may include six forward electric propulsion systems operably connected to a tilting propeller and four tail electric propulsion systems operably connected to a lift propeller. In some embodiments, some or all of the tail electric propulsion systems may be operably connected to the tilting propeller.

[0080] like Figure 9E As shown, in some embodiments, the aircraft may have a flying wing configuration, such as a tailless fixed-wing aircraft without a fixed fuselage. In some embodiments, the aircraft may have a flying wing configuration with the fuselage integrated into the wing. In some embodiments, when the tiltrotor operates in a lift configuration, the tiltrotor may rotate in a plane above the aircraft fuselage.

[0081] As disclosed herein, the front electric propulsion system and the rear electric propulsion system can be of clockwise (CW) or counterclockwise (CCW) type. Some embodiments may include various front 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 within the rear electric propulsion system. In some embodiments, each electric propulsion system may be fixed to be of clockwise (CW) or counterclockwise (CCW) type, while in other embodiments, one or more electric propulsion systems may vary between clockwise (CW) rotation and counterclockwise (CCW) rotation.

[0082] Figure 10 A functional block diagram of an exemplary control system 1000 for an aircraft consistent with the disclosed embodiments is shown. 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., a computer-readable medium, a non-transitory 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 the flight control system of an aircraft (e.g., Figure 6 (As part of FCS 612), and can be configured to repeatedly execute a single step or sequence until the desired or commanded result is achieved. It should be understood that, for ease of description, Figure 10Many conventional functions of the control system are not shown. System 1000 further includes one or more storage media storing models, functions, tables, and / or any information used to execute the disclosed processes. As further described below, any or each of the boxes indicating command models 1004, 1006, 1008, 1010, feedback 1012, 1016, 1018, 1022, feedforward 1014, 1020, outer loop assignment 1024, 1026, inner loop control law 1028, control assignment 1029, and DCPS 1033 may represent or contain modules, scripts, functions, applications, and / or programs executed by the processor and / or microprocessor of system 1000. It should be understood that Figure 10 The complexity and interconnectivity of the functional block diagrams make it impossible, or at least impractical, for human users to implement the method effectively, especially when considering the implementation of these functions while the aircraft is in flight (including takeoff or landing) (e.g., in real time). Unless otherwise stated herein, “expectations” regarding aircraft commands, aircraft parameters, or other aircraft characteristics can refer to characteristics that are inputs (e.g., via pilot input devices, system modules, etc.), requests, and / or otherwise presented as part of the process for controlling the computationally relevant operations of the aircraft.

[0083] In some embodiments, the control system 1000 may be configured based on one or more flight control laws. Flight control laws may include a set of algorithms, models, and / or rules configured to manage the behavior of an aircraft (e.g., control or influence one or more actuators of the aircraft) in response to one or more pilot inputs and external factors. In some embodiments, flight control laws may be configured to achieve at least one of desired flight characteristics, stability, or performance. For example, flight control laws 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, aircraft dynamics (e.g., disturbances, such as turbulence, gusts, etc.), or changes in flight conditions (e.g., altitude, airspeed, angle of attack, torque of one or more propellers).

[0084] System 1000 can detect one or more inputs, for example, from a pilot input device configured to receive at least one pilot input and generate or influence signals. Pilot inputs can be generated and / or received from input devices or mechanisms of the aircraft, such as buttons, switches, joysticks, sliders, actuators, or any other means configured to generate or influence signals based on physical actions from the pilot. For example, a pilot input device may include one or more of a right actuator (e.g., left / right movement 1002a and / or forward / backward movement 1002e), a left actuator (e.g., left / right movement 1002c and / or forward / backward movement 1002g), and / or a left actuator switch 1002f. In some embodiments, the pilot input device may include an interface to an autopilot system (e.g., a display screen, switch, button, joystick, and / or other interface). Optionally, the system 1000 may further detect input from the autopilot system, such as autopilot roll command 1002b, autopilot climb command 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 and / or left controllers, signals received from switches on the controllers (e.g., response type change commands, trim inputs, reference inputs, backup control inputs, etc.), measurements of 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., ground taxiing, takeoff, airborne). For example, the right controller L / R 1002a may include the lateral position and / or rate of the right controller (e.g., a controller positioned to the right of another controller and / or a controller positioned to the right of the pilot area), the autopilot roll command 1002b may include a roll signal received in autopilot mode, the left controller L / R 1002c may include the lateral position and / or rate of the left controller (e.g., a controller positioned to the left of another controller and / or a controller positioned to the left of the pilot area), the autopilot climb command 1002d may include a climb signal received in autopilot mode, the right controller F / A 1002e may include the longitudinal position and / or rate of the right controller, the left controller switch 1002f may include a signal from a switch for enabling or disabling the automatic transition function 1003, and the left controller F / A 1002g may include the longitudinal position and / or rate of the left controller.

[0086] At least one (e.g., each) input may contain data as listed above (e.g., signals from switches, measurements of aircraft status, aircraft modes, etc.). Actuator status may include actuator hardware limitations, such as travel limits, speed limits, response time limits, etc., and may include actuator health indicators that can indicate deterioration in actuator performance, potentially limiting a given actuator's ability to fulfill actuator commands. Actuator status can be used to determine the limits (e.g., minimum / maximum) of individual actuator commands. Battery status may correspond to the remaining energy of the aircraft's battery pack, which can be monitored when control allocation 1029 considers optimizing the battery pack's energy status. Aerodynamic parameters may be parameters derived from aerodynamic and acoustic modeling and may be based on actuator Jacobian matrices (e.g., describing how changes in the actuators (e.g., in response to actuator / actuator commands) affect the overall motion of the aircraft) and actuator status. At least one (e.g., each) input received from the controller may indicate a corresponding adjustment to the aircraft's heading or power output.

[0087] Command models 1004, 1006, 1008, and 1010 can be configured to determine the shape of an ideal aircraft response (e.g., aggressiveness, rate of turn, damping, overshoot, etc.). For example, at least one (e.g., each) of command models 1004, 1006, 1008, and 1010 can be configured to receive and interpret at least one of inputs 1002a, 1002b, 1002c, 1002d, 1002e, 1002f, and 1002g, and, in response, use an integrator (not shown) to calculate corresponding changes in the aircraft's direction, heading, and propulsion, or combinations thereof. In some embodiments, the right control L / R 1002a and autopilot roll command 1002b can be fed into the turning rate command model 1004, the left control L / R 1002c can be fed into the lateral speed command model 1006, the autopilot climb command 1002d and right control F / A 1002e can be fed into the climb command model 1008, and the left control F / A 1002g can be fed into the forward speed command model 1010. In some embodiments, the output from the automatic transition function 1003 can be fed into 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 control switch 1002f, the automatic transition function 1003 can automatically determine at least one of the climb signal or the forward speed signal for transmission to at least one of the climb command model 1008 or the forward speed command model 1010.

[0088] Turn rate command model 1004 can be configured to output desired position and / or turn rate commands, and can also be configured to calculate the assumed desired heading of the aircraft when the controller is brought back to the center position (i.e., in a slack position). Lateral speed command model 1006 can be configured to output desired position and / or lateral speed commands. Climb command model 1008 can be configured to output at least one of desired altitude, vertical speed, or vertical acceleration commands. Forward speed command model 1010 can be configured to output at least one of desired position, longitudinal speed, or longitudinal acceleration commands. In some embodiments, one or more of the command models can be configured to output acceleration generated in response to changes in the speed command. For example, climb command model 1008 can be configured to output vertical acceleration generated in response to changes in the vertical speed command.

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

[0090] At least one of the feedbacks 1012, 1016, 1018, and 1022 (e.g., each) may receive one or more desired changes (e.g., desired position, velocity, and / or acceleration) from command models 1004, 1006, 1008, and 1010, as well as data indicating aircraft dynamics 1030 received from aircraft sensing 1031 as input. For example, the sensed aircraft dynamics 1030 may include a representation of the aircraft's physical and / or natural dynamics, and the aircraft dynamics sensing 1031 sensor measurements may capture how the aircraft moves in response to pilot input, propulsion system output, or environmental conditions. Additionally or alternatively, the data received from aircraft sensing 1031 may include error signals generated by one or more processors based on external disturbances (e.g., speed disturbances caused by gusts). In some embodiments, the feedbacks 1012, 1016, 1018, and 1022 may be configured to generate feedback forces (e.g., at actuators) based on the received error signals. For example, feedbacks 1012, 1016, 1018, and 1022 can generate feedback forces to counteract the effects of external disturbances. Alternatively, feedbacks 1012, 1016, 1018, and 1022 can be configured to generate feedback forces based on modeling errors. For example, if incorrect aircraft mass is input into feedforward 1014 or 1020, the aircraft's acceleration may change faster or slower than expected. Based on determining the difference between the desired acceleration and the measured acceleration, one or more processors (e.g., included in aircraft sensing 1031) can generate an error signal, which can be looped through feedbacks 1012, 1016, 1018, or 1022 to determine the additional force required to correct the error.

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

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

[0093] In some embodiments, feedback 1012, 1016, 1018, 1022 may determine one or more feedback forces based on actuator state information received from one or more sensors (e.g., included in aircraft sensing 1031). For example, in response to actuator state information indicating the presence of an actuator malfunction, one or more processors of system 1000 may update one or more processes of system 1000 and determine alternative commands for achieving the 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 response to) a state change (e.g., malfunction) of an actuator (or other aircraft component, such as an engine or battery, for other examples). Alternative commands (e.g., yaw, pitch, roll, thrust, or torque) may be determined based on the adjustment. Additionally or alternatively, in response to actuator state information indicating that one or more actuators are at their maximum values, one or more processors of system 1000 may update one or more processes of system 1000 (e.g., as described above) and determine alternative commands for achieving the desired change.

[0094] The total desired force can be calculated based on the outputs of feedbacks 1012, 1016, 1018, 1022 and feedforwards 1014, 1020. For example, one or more processors of system 1000 can calculate the desired turning rate force by summing the outputs of feedback 1012 and feedforward 1014. Alternatively, one or more processors of system 1000 can calculate the desired lateral force by summing the outputs of feedback 1016 and feedforward 1014. Alternatively, one or more processors of system 1000 can calculate the desired vertical force by summing the outputs of feedback 1018 and feedforward 1020. Alternatively, one or more processors of system 1000 can 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 assignment 1024 and the longitudinal outer loop assignment 1026 (e.g., each) can be configured to receive one or more desired forces and data (e.g., airspeed, aircraft orientation, aircraft load factor, measured acceleration, aircraft mass and inertia, indications of working / failed actuators, air density, altitude, aircraft mode, whether the aircraft is in the air or on the ground, etc.) received from the aircraft sensor 1031 as input. Based on the input, the outer loop assignments 1024 and 1026 can be configured to command roll, command yaw, command pitch, request thrust, or output a combination of different commands / requests to achieve one or more desired forces.

[0096] The lateral / directional outer loop assignment 1024 can receive desired turn rate force and / or desired lateral force as input and can command roll or command yaw. In some embodiments, the lateral / directional outer loop assignment 1024 can determine the output based on a determined flight mode. The flight mode can be determined using pilot input (e.g., a selected mode on the controller) and / or sensed aircraft information (e.g., airspeed). For example, the lateral / directional outer loop assignment 1024 can determine the aircraft's flight mode using at least one of a determined (e.g., sensed or measured) airspeed or input received at a pilot controller button (e.g., an input instructing the aircraft to fly according to a specific flight mode). In some embodiments, the lateral / directional outer loop assignment 1024 can be configured to prioritize pilot controller button input over measured airspeed when determining the flight mode (e.g., pilot controller buttons are associated with a higher weight or priority than measured airspeed). In some embodiments, the lateral / orientation outer loop allocation 1024 can be configured to combine (e.g., using weighted summation) determined airspeed and pilot control button inputs to determine the aircraft's flight mode. In hover flight mode, the lateral / orientation outer loop allocation 1024 can utilize roll commands (e.g., roll angle, roll rate) to achieve the desired lateral force and can utilize yaw commands to achieve the desired turn rate force. In some embodiments, such as in hover flight mode, the aircraft can be configured not to accelerate outside a predetermined hover envelope (e.g., hover speed range). In forward flight mode (e.g., level flight), the lateral / orientation outer loop allocation 1024 can utilize yaw commands to achieve the desired lateral force and can utilize roll commands to achieve the desired turn rate force. In forward flight mode, the lateral / orientation outer loop allocation 1024 can be configured to determine the output based on sensed airspeed. During the transition between hover flight mode and forward flight mode, the lateral / orientation outer loop allocation 1024 can use a combination of roll and yaw commands to achieve the desired force.

[0097] The longitudinal outer loop assignment 1026 can receive a desired vertical force and / or a desired longitudinal force as input, and can output at least one of a pitch command (e.g., pitch angle) or a thrust vectoring request. The thrust vectoring request can include longitudinal thrust (e.g., a mixture of nacelle tilt propeller thrust and nose propeller thrust) and vertical thrust (i.e., a combination of forward and backward thrust). In some embodiments, the longitudinal outer loop assignment 1026 can determine the output based on a determined flight mode. For example, in a hovering flight mode, the longitudinal outer loop assignment 1026 can achieve the desired longitudinal force by reducing the pitch attitude and by using longitudinal thrust, and can achieve the desired vertical force with vertical thrust. In a forward flight mode, the longitudinal outer loop assignment 1026 can achieve the desired longitudinal force with longitudinal thrust (e.g., nose propeller thrust). In a cruise flight mode, the longitudinal outer loop assignment 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 can be configured to determine a torque command based on at least one of a roll command, yaw command, or pitch command from the lateral / orientation outer-loop assignment 1024 or the longitudinal outer-loop assignment 1026. In some embodiments, the inner-loop control law 1028 may depend on sensed aircraft dynamics (e.g., from aircraft sensing 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. Alternatively or additionally, the inner-loop control law 1028 may take into account the period of natural patterns (e.g., undulation patterns) affecting the pitch axis and may appropriately control the aircraft to compensate for such natural patterns. In some embodiments, the inner-loop control law 1028 may depend on aircraft inertia.

[0099] The inner-loop control law 1028 may use one or more stored dynamic models reflecting the aircraft's motion characteristics (e.g., the aircraft's aerodynamic damping and / or inertia) to determine the torque command. 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 aircraft's motion characteristics and determine one or more torques that will cause the aircraft to achieve the commanded roll, yaw, and / or pitch. Some embodiments may include determining (e.g., via the inner-loop control law 1028 or other components) the torque command based on at least one received command (e.g., roll command, yaw command, and / or pitch command) and a determined (e.g., measured) aircraft state. For example, the torque command may be determined using the difference between the commanded aircraft state and the measured aircraft state. As a further example, the torque command may be determined using the difference between the commanded roll angle and the measured roll angle. As described below, control assignment 1029 may control the aircraft (e.g., via flight elements) based on the determined torque command. For example, control allocation 1029 can control one or more electric propulsion systems of an aircraft (e.g., Figure 6 The electric propulsion system 602 shown (e.g., transmitting one or more commands to one or more electric propulsion systems), said one or more electric propulsion systems including a tilting actuator, an electric engine, and / or a propeller. Control assignment 1029 can further control one or more control surfaces of the aircraft (e.g., control surfaces, such as...). Figure 7 The flaperon 712 and rudder 714 shown include flaperons, rudders, ailerons, spoilers, rudders, and / or elevators. Aircraft dynamics 1030 represents the control of various flight components (e.g., electric propulsion systems and / or control surfaces) and the corresponding effects on the flight components and aircraft dynamics.

[0100] Although Figure 10 The illustrated embodiment includes both inner-loop control law 1028 and outer-loop assignments 1024, 1026; however, in some embodiments, the flight control system may not include outer-loop assignments 1024, 1026. Therefore, pilot control inputs can create roll, yaw, pitch, and / or thrust commands. For example, the right control can control roll and pitch, and the left control and / or pedals can control yaw and thrust.

[0101] Control allocation 1029 may accept one or more of the following as inputs: force or torque commands, data received from one or more aircraft sensors, envelope protection limits, scheduling parameters, and optimizer parameters. Control allocation 1029 may be configured to determine actuator commands (e.g., thrust, torque, and / or propeller speed for an electric propulsion unit) based on the inputs by minimizing an objective function (e.g., solving an optimization problem, e.g., continuously), the objective function comprising one or more primary objectives, such as satisfying (e.g., responding to, conforming to, resolving, or providing output based on) the aircraft forces and torques of the commands, and one or more secondary objectives, which may include minimizing acoustic noise and / or optimizing battery pack usage.

[0102] In some embodiments, control assignment 1029 may be configured to calculate limits for individual actuator commands based on actuator states and envelope protection limits. Envelope protection limits, as used herein, may include one or more limits within 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 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, control assignment 1029 may use envelope protection limits to automatically adjust one or more control surfaces or control settings. This prevents undesirable conditions such as stall, structural strain, or failure of the aircraft. During normal operation, the minimum command limit for a given actuator may include the maximum of the following: a hardware-based minimum limit and a minimum flight envelope limit; and the maximum command limit for a given actuator may include the minimum of the following: a hardware-based maximum limit and a maximum flight envelope limit. In the event of an actuator failure, the command limit of the failed actuator corresponds to the failure mode.

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

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

[0105] As used herein, the term "actuator" refers to 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” refers to any combination of electrically connected batteries (i.e., battery cells) and may contain multiple batteries arranged in series, in parallel, or in a combination of series and parallel.

[0107] In some embodiments, control allocation 1029 may include engine thermal management functions to manage (e.g., optimize) 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 higher temperatures relative to other electric propulsion units. (See below) Figure 20 and Figure 21 The description provides a detailed description related to engine thermal management. In some embodiments, control allocation 1029 may perform both energy optimization and engine thermal management. In some embodiments, control allocation 1029 may prioritize engine thermal management over energy optimization.

[0108] Control assignment 1029 sends commands to one or more flight elements to control the aircraft. The flight elements will move according to the control commands. Various sensing systems and associated sensors, as part of aircraft dynamic sensing 1031, can detect the movement of the flight elements and / or the dynamics of the aircraft, and provide information to feedback 1012, 1016, 1018, 1022, outer loop assignment 1024, 1026, inner loop control law 1028, and control assignment 1029 to be incorporated into flight control.

[0109] As described above, the aircraft sensing 1031 may include one or more sensors to detect aircraft dynamics. For example, the aircraft sensing 1031 may capture how the aircraft moves in response to pilot input, propulsion system output, or environmental conditions. Alternatively or additionally, the aircraft sensing 1031 may detect errors in the aircraft response based on external disturbances (e.g., speed disturbances caused by gusts). Furthermore, the aircraft sensing 1031 may include one or more sensors for detecting propeller speed, such as magnetic sensors (e.g., Hall effect or inductive sensors) or optical sensors (e.g., tachometers) for detecting the rotor speed of the aircraft engine (and thus the propeller speed). The aircraft sensing 1031 may include one or more sensors to detect nacelle tilt angles (e.g., lift configurations (e.g., ...). Figure 2 ) and forward thrust configuration (e.g., Figure 1 The one or more sensors, such as magnetic sensors (e.g., Hall effect or inductive sensors), may include one or more sensors configured to detect engine torque and / or thrust, such as one or more current or voltage sensors, strain gauges, force gauges, and / or propeller vibration sensors (e.g., accelerometers).

[0110] Aircraft sensing 1031 may include one or more sensors configured to detect aircraft dynamics, such as acceleration and / or pitch direction sensors (e.g., accelerometers, 3-axis accelerometers, gyroscopes, 3-axis gyroscopes, and / or tilt position sensors for determining engine angles) and airspeed sensors (e.g., pitot tube sensors). Aircraft sensing 1031 may further include one or more inertial measurement units (IMUs) to determine the aircraft state based on these measurements. Aircraft state may refer to forces experienced by the aircraft, the aircraft's orientation, the aircraft's position (e.g., altitude), and / or the aircraft's movement. For example, 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 direction across one or both axes), the aircraft's speed, the aircraft's angular rate (e.g., roll, pitch, and / or yaw rate), and / or the aircraft's acceleration (e.g., longitudinal, lateral, and / or vertical acceleration), or any physical characteristic of the aircraft or one of its components. In some embodiments, the aircraft sensing 1031 may include an inertial navigation system (INS) and / or an air data and / or attitude heading reference system (ADAHRS). The INS and / or ADHRS may include one or more inertial measurement units (IMUs) and corresponding sensors (e.g., accelerometers, gyroscopes, three-axis gyroscopes, and / or three-axis accelerometers). In some embodiments, the INS and / or ADHRS may filter and / or otherwise process sensor measurements to determine the aircraft's state (e.g., acceleration or angular rate). For example, in some embodiments, the INS and / or ADHRS may determine the angular rate based on gyroscope measurements and may determine the acceleration based on measurements from the accelerometer.

[0111] DCPS 1033 may receive input from control assignment 1029 and aircraft sensing 1031. For example, DCPS 1033 may receive at least one or more torque commands, one or more limits (e.g., envelope protection limits, engine torque rate limits, HV channel / HV bus protection torque limits, etc.), or measured aircraft dynamics (e.g., measured revolutions per minute (RPM), measured voltage, etc.). DCPS 1033 may be configured to modify one or more torque commands based on the input to dynamically generate one or more modified torque commands (e.g., dynamically varying torque commands). System 1000 may be configured to send one or more dynamically modified torque commands to one or more engines of the aircraft. The disclosed embodiments may improve engine response to enhance aircraft stability and safety.

[0112] Figure 11A block diagram depicts an exemplary flight control system 1100 including energy optimization functions according to various embodiments. It should be understood that human users effectively implement... Figure 11 The steps of the exemplary methods described herein would be impossible, or at least impractical, especially when considering that these functions are frequently (e.g., continuously, constantly) implemented during aircraft flight (including takeoff or landing), and / or dynamically implemented based on (e.g., in response to) received signals (e.g., aircraft sensors, pilot input devices). Generally, it can be understood that... Figure 11 Any / all steps of the exemplary method may be executed or performed by at least one processor (e.g., FCS, System 1000), for example, according to one or more instructions stored on a computer-readable medium (e.g., a non-transitory 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 torques based on one or more updated initial reference commands and calculate allocations (e.g., as per the context of...). Figure 10 The control assignment described herein is to satisfy (e.g., response, compliance, resolution, based on) recalculated forces and / or moments, and to minimize the differences from one or more updated initial reference commands. It should be further understood that, although Figure 11 An example of energy optimization related to torque is depicted, but Figure 11 It may be associated, or otherwise, with any parameters (such as speed, power, and / or current) related to aircraft actuators.

[0113] At 1110, the energy optimization function can update one or more initial reference commands 1102 (e.g., reference torque command). In some embodiments, the energy optimization function may update one or more initial reference commands 1102 based on battery information 1104 (e.g., from the BMS, such as available energy (e.g., usable energy)), (e.g., the remaining discharge time of one or more battery packs, one or more crosslinks, one or more independent high-voltage buses, one or more high-voltage channels), state of energy (SOE), state of charge (SOC), state of power (SOP), state of health (SOH), fault conditions (e.g., short circuit or overcurrent, whether the battery is activated / operating), or the 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 with low available energy (e.g., relative to other battery packs, relative to the average available energy of all available (e.g., activated, operating) battery packs), the energy optimization function may perform allocation functions (e.g., solving for...) as needed. Figure 10 Before performing the allocation function (as described in the control allocation 1029), one or more first initial reference commands of one or more initial reference commands 1102 associated with one or more first rotors are updated to a first range. Alternatively, if one or more second rotors are connected to a second battery pack with high available energy (e.g., relative to other battery packs, relative to the average available energy of all available (e.g., active, functioning) 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 a second range above 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 state differences between at least two battery packs (e.g., between each battery pack) based on battery information 1104 received from one or more battery management systems associated with multiple battery packs. In some embodiments, the energy optimization function may be configured to calculate one or more updated initial reference torque commands as follows: in: The usable energy (e.g., available energy) discussed herein can be determined by at least one BMS or flight control system based on battery information (e.g., measured information), such as (e.g., one or more battery packs, one or more crosslinks, one or more independent high-voltage buses, one or more high-voltage channels) remaining discharge time, state of energy (SOE), state of charge (SOC), state of power (SOP), state of health (SOH), fault condition (e.g., short circuit or overcurrent, whether the battery is activated / functioning), or the state of temperature (SOT) of at least one (e.g., each) battery pack.

[0114] In some embodiments, the energy optimization function may determine one or more initial reference commands 1102. Alternatively, the energy optimization function may receive one or more initial reference commands 1102 from a flight control law (e.g., as discussed with respect to system 1000). For example, the flight control system may use the 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 of the following based on one or more lookup tables: reference commands (e.g., initial reference torque command, initial reference power command, initial reference speed command, initial reference current command), reference actuator position, reference force, reference torque, and / or one or more Jacobian expressions (e.g., Jacobian matrix). For example, the lookup model may refer to one or more lookup tables containing scheduling parameters and using one or more current aircraft states (e.g., […]). Figure 10The aircraft dynamics, environmental conditions, flight status (e.g., altitude, airspeed, acceleration, angle of attack, torque of one or more propellers, tilt angle and / or direction of one or more propellers, whether powered lift is activated, acceleration, load factor of driven flap and aileron scheduling, control surface position / movement, flight phase, speed of one or more propellers, and / or engine torque of one or more engines), and the status and / or health condition 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, etc.) are used to determine reference values ​​and Jacobians. Scheduling parameters may include one or more of the following: EPU scheduling (e.g., coordinating rotor / propeller speed and direction during different phases of flight), battery management (e.g., varying the power drawn from different battery packs to ensure optimal performance and lifespan), control surface movement (e.g., coordinating the movement of flaps, elevators, rudders, or any other control surfaces during different phases of flight), power allocation (e.g., ensuring that critical systems receive priority power during any high-demand conditions), timing of inputs for managing 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 that describes how to manage or coordinate different components of the aircraft for various aircraft conditions to ensure optimal performance, safety, and efficiency.

[0115] In some embodiments, the energy optimization function can be configured to limit one or more initial reference commands. For example, the energy optimization function can automatically limit 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 can limit the initial rotor reference torque within an envelope protection range (e.g., below the maximum envelope protection limit and above the minimum envelope protection limit). In some embodiments, the energy optimization function can receive one or more limited initial reference commands (e.g., limited by the flight control system using flight control laws).

[0116] At 1120, the flight control system can recalculate (e.g., adjust) one or more force or torque (FM) commands. For example, the flight control system can recalculate (e.g., adjust) one or more initial force or torque commands 1122. Figure 10 The control assigns input 1029. And based on one or more outputs of 1110 (e.g., one or more variations (e.g., incremental changes) of the initial reference command 1106, the difference between the updated reference command and the initial reference command), The flight control system may recalculate one or more force or torque commands as follows: In some embodiments, the flight control system may recalculate one or more force or torque commands as follows: in: At 1130, the flight control system can perform control assignment (e.g., solve the assignment, by) based on the assignment input 1132, the updated initial reference command 1108, and the recalculated force / torque command 1124. Figure 10 The control allocation 1029 is executed in the steps of execution). In some embodiments, due to (e.g., based on) the execution of step 1130, the flight control system may output one or more actuator commands 1134 (e.g., Figure 10 The control allocation of the 1029 output, (etc.). For example, based on the received allocation input 1132 (e.g., Figure 10 The output of the inner-loop control law 1028, the output of the outer-loop allocation 1026 and / or the input of the control allocation 1029, data received from one or more aircraft sensors, envelope protection limits, scheduling parameters and / or optimizer parameters), flight control system (e.g., Figure 10 The control allocation 1029 can be solved such that the recalculated force and / or torque command 1124 is satisfied (e.g., fully within the threshold margin) and the difference between the updated initial reference command 1108 and one or more output actuator commands 1134 is minimized.

[0117] In some embodiments, the energy optimization function may be considered a secondary objective relative to at least one other objective or function (e.g., implemented by the flight control computer and / or by 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 discussed below. In some embodiments, the electric engine timescale may be shorter than the battery energy optimization timescale. For example, the FCC or another device implementing the different functions disclosed herein may optimize electric engine temperature parameters over a shorter timescale than the timescale 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 outputs that are opposite to each other, the flight control system may associate a reference value associated with the engine thermal management function with a greater weight than a 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 flight phase, the temperature and / or remaining energy of one or more aircraft components (e.g., engines, batteries).

[0118] 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 between different battery components that may have different time constants and temperature limits using a time-based temperature metric.

[0119] Figure 12 This is a flowchart illustrating an exemplary method 1200 for energy optimization of a particular machine (e.g., an aircraft) according to some embodiments of the present disclosure, thereby improving the aircraft's technology (e.g., aircraft safety, energy efficiency, range capability, payload capacity, structural integrity). The steps of method 1200 can be, for example, in conjunction with... Figures 1 to 8 , Figures 9A to 9E , Figure 10 , Figure 11 , Figures 13A to 13D , Figures 14 to 2 Performed or otherwise used on any aircraft component associated with any of the figures in 5 Figure 10 The method may be executed by system 1000, or by any flight control computer (e.g., a computer-implemented method) or flight control system. For example, the aircraft's flight control computer may be configured to execute one or more steps of method 1200.

[0120] The steps of method 1200 can be activated or adjusted as needed based on aircraft flight conditions, maneuvering, or operational requirements. It should be understood that the illustrated method 1200 can be modified to alter the order of steps and include additional steps. It should also be understood that the complexity of method 1200 would make it impossible, or at least highly impractical, for a human user to implement the method effectively, especially when considering the implementation of these functions while the aircraft is in flight (including takeoff or landing) (e.g., in real time), and when the energy used by the aircraft and the battery status are constantly changing. Furthermore, energy optimization (e.g., using method 1200) improves aircraft safety and range capabilities (e.g., how much energy is available to power the aircraft) and the lifespan of different components (e.g., batteries) without requiring any direct involvement from the user (e.g., pilot). In some embodiments, the aircraft of system 1000 may include an electric aircraft comprising multiple actuators, each actuator comprising multiple electric propulsion units (EPUs) and multiple battery packs powering the multiple EPUs.

[0121] In step 1202, system 1000 may determine one or more desired commands for an electric aircraft. For example, the one or more desired commands may include at least one of the following: one or more force or torque commands (e.g., Figure 10 The inputs to the control assignment 1029), desired position and / or turning rate commands (e.g., Figure 10 The output of the turning rate command model 1004), the desired position and / or lateral speed command (e.g., Figure 10 The output of the lateral velocity command model 1006), one or more of the desired height, vertical velocity, or vertical acceleration commands (e.g., Figure 10 The output of the climb command model 1008), the desired position, longitudinal velocity, or longitudinal acceleration command (e.g., Figure 10 The output of the forward speed command model 1010), and alternative commands used to achieve the desired change (e.g., based on such as...). Figure 10 The feedback discussed in the article), one or more total expectancy (e.g., based on Figure 10 The outputs of feedback 1012, 1016, 1018, 1022 and feedforward 1014, 1020 are calculated), and one or more torque commands (e.g., by...) Figure 10The inner loop control law 1028 is determined), linear command, angular acceleration command, pitch command, roll command, yaw command, or any command associated with the desired motion of the aircraft (e.g., the desired motion of the aircraft's pilot, the desired motion of the flight control system, or the desired motion of the autopilot system) in the form of one or more signals (e.g., pilot input from the pilot input device, input from the autopilot system, signals generated by the flight control system or computer).

[0122] In step 1204, 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 states, ideal actuator / actuator position / velocity / torque / current / power) to achieve one or more desired commands. For example, system 1000 may be configured to continuously monitor the aircraft's condition or state (e.g., using a feedback control system) and may generate or adjust control commands (e.g., adjusting force / torque commands, adjusting aircraft components, such as actuators, actuators, propulsion units, control surfaces, etc.) to correct any deviations from the reference command. In some embodiments, system 1000 may determine at least one reference command based on one or more aircraft conditions. Aircraft conditions may include one or more of the following: (e.g., Figure 10 The system determines the status and / or health of at least one aircraft component (e.g., altitude, airspeed, angle of attack, torque of one or more propellers, roll angle of one or more propellers, whether powered lift is activated, acceleration, load factor of actuated flaps and ailerons, flight phase, speed of one or more propellers, engine torque of one or more engines), and / or the status of at least one (e.g., each) aircraft component (e.g., availability of batteries, propulsion units, high-voltage buses and / or channels, etc.). In some embodiments, determining at least one reference command may include using multiple predetermined lookup tables to determine at least one initial (e.g., predetermined, stored) reference command based on the current aircraft status. In some embodiments, the system 1000 may determine at least one reference command for at least one (e.g., each) actuator, actuator, or both. In some embodiments, the system 1000 may determine at least one reference command for at least one (e.g., each) force, torque, or both. In some embodiments, 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.

[0123] In step 1206, system 1000 can monitor the state of energy (SOE) of multiple battery packs. SOE can be represented by and / or include one or more battery parameters (e.g., related to consumed energy, remaining energy, available energy, usable energy, and / or energy to be consumed) indicating the battery pack's measurement results, predictions, and / or capabilities. In some embodiments, system 1000 can receive battery parameters associated with at least one (e.g., each) battery pack from one or more BMSs associated with the multiple battery packs. Battery parameters can include one or more of the following: usable energy of at least one (e.g., each) battery pack, remaining discharge time (e.g., one or more batteries, one or more crosslinks, one or more independent high-voltage buses, one or more high-voltage channels), SOE, SOC, SOP, SOH, fault condition (e.g., short circuit or overcurrent, whether the battery is activated / functioning / malfunctioning / faulty), or SOT. In some embodiments, system 1000 can calculate at least one of usable energy or remaining discharge time based on other battery information. In some embodiments, battery information may include differences in power consumption between the engines of multiple EPUs (e.g., EPU engines). In some embodiments, system 1000 may be communicatively coupled (e.g., via physical connection, such as a bus and / or channel system, or via a digital communication interface) to one or more battery management systems (“BMS”) of an aircraft. In some embodiments, the aircraft may contain a single BMS configured to manage all battery packs on the aircraft. In some embodiments, at least one (e.g., each) battery pack may contain its own BMS.

[0124] In some embodiments, system 1000 may also monitor the status of at least one (e.g., each) of 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) of a plurality of EPUs (e.g., via a digital communication interface, via an inverter of the EPU). The EPU information may include one or more of the following: at least one temperature of one or more components (e.g., engine, propeller) of at least one (e.g., each) EPU; one or more times associated with at least one temperature of one or more components of at least one (e.g., each) EPU (e.g., the remaining time the engine has at the current torque rating, such as continuous torque limit, etc.). Figure 22 2204), engine takeoff torque limit (e.g., Figure 22 2205), engine emergency torque limit (e.g., Figure 222206), engine overheat torque limit (e.g., Figure 22 (2207), and the expected range at maximum takeoff torque (e.g., Figure 22 (2208), or the expected range at maximum emergency torque (e.g., Figure 22 (2209) or fault conditions (e.g., whether the EPU and / or components of the EPU are activated and respond 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 at one or more future times (e.g., points in time), as discussed further below. In some embodiments, 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 depleted (or nearly depleted, e.g., depleted below a specific threshold) at the same or near-the-same time (e.g., within tolerance at the same time).

[0125] In step 1208, system 1000 may adjust at least one reference command (e.g., an initial reference command) based on the monitored state of energy of multiple battery packs. In some embodiments, adjusting at least one reference command may include limiting at least one reference command using one or more command limits (e.g., envelope protection limits, minimum and / or maximum limits associated with actuators / actuators, and at least one command parameter (e.g., position, speed, torque, power, current) based on dynamic aircraft conditions and / or predetermined maximum / minimum structural limits) such that the limited reference command is within command limits (e.g., below or above command limits).

[0126] 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) of a plurality of high-voltage buses / channels. In some embodiments, the one or more estimated reference commands may include at least one of the following: 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 high-voltage architecture of the aircraft 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 engines / rotors and HV channels / buses, between battery packs and engines / rotors, between battery packs). In some embodiments, the connection matrix may be stored in a flight control law. In some embodiments, the system 1000 may dynamically update the connection matrix based on dynamic updates to the connections (e.g., based on identifying faults in one or more components connected to the high-voltage architecture, based on faulty connections).

[0127] In some embodiments, adjusting at least one reference command may include 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 balance 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 per high-voltage bus / channel based on the average of the total commands (e.g., total power command, total current command, total position command, total balance command, total torque command, and / or total speed command) of the high-voltage bus / channel. For example, system 1000 may determine the power ratio per high-voltage bus as follows: in: 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. Alternatively or additionally, 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: in: In some embodiments, adjusting at least one reference command may include updating one or more reference commands to optimize the remaining discharge time on all high-voltage buses / channels. For example, system 1000 may adjust at least one reference command (e.g., initial reference command, limiting reference command) based on one or more determined ratios (e.g., power ratio, energy ratio, current ratio, torque ratio, speed ratio) and a connection matrix. In some embodiments, system 1000 may adjust at least one reference command based on the relationship between at least two ratios. For example, system 1000 may adjust at least one reference command based on the relationship between one or more determined power ratios and one or more determined energy ratios (e.g., ratios). In some embodiments, system 1000 may determine (e.g., optimize) the remaining discharge time of at least one (e.g., each) high-voltage bus / channel based on at least one reference command, one or more determined ratios, or one or more of the connection matrix. For example, system 1000 may update one or more reference commands to balance (e.g., equalize to within a threshold) the remaining discharge time on all buses / channels, as follows: in: 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 of at least one (e.g., each) high-voltage bus / channel is the same. Alternatively, determining (e.g., optimizing) the remaining discharge time may include adjusting at least one reference command such that the remaining discharge time of at least one (e.g., each) high-voltage bus / channel is within a predetermined range.

[0128] In some embodiments, when the monitored state of energy indicates that the first engine has lower power consumption than the second engine, the system 1000 may adjust a first reference command associated with the first engine and a second reference command associated with the second engine. For example, the system 1000 may increase the torque or power command associated with the first reference command relative to the torque or power command associated with the second reference command.

[0129] In some embodiments, the system 1000 may further adjust at least one reference command based on the monitoring status of at least one (e.g., each) of a plurality of electric propulsion units. For example, based on received EPU information, the system 1000 may adjust at least one reference command to 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 components thereof, such as propellers or engines).

[0130] In some embodiments, system 1000 may further adjust at least one reference command based on predictions of power usage (e.g., power required or used by one or more components of an aircraft). For example, system 1000 may use simulated data (e.g., stored simulated data) or historical data to perform power usage predictions based on one or more aircraft conditions. In some embodiments, the simulated data and / or historical data may indicate the power usage of at least one (e.g., each) engine associated with different aircraft conditions. In some embodiments, the simulated data or historical data may include one or more of flight test data, weather data, route data, expected discharge data, range data, or battery health data.

[0131] In some embodiments, system 1000 may limit at least one adjusted reference command. For example, system 1000 may use one or more command limits (e.g., envelope protection limits, minimum and / or maximum limits associated with actuators / actuators, and command parameters (e.g., position, speed, torque, power, current) based on dynamic aircraft conditions and / or predetermined maximum / minimum structural limits) to limit at least one adjusted reference command such that the limited adjusted reference command is within (e.g., below, above) the command limit. In some embodiments, the adjusted at least one reference command in steps 1210 and 1212 may refer to the limited adjusted at least one reference command.

[0132] In step 1210, system 1000 may generate control commands for multiple actuators based on at least one adjusted reference command. In some embodiments, generating control commands for multiple actuators may include calculating the difference between force and torque based on at least one adjusted reference command. In some embodiments, generating control commands may include recalculating the force or torque based on at least one adjusted reference command (e.g., as...). Figure 11 (as described above). For example, system 1000 can recalculate force and / or torque by determining changes in force and / or torque as follows: in: In some embodiments, generating control commands for multiple actuators may include calculating actuator limits based on at least one adjusted reference command. In some embodiments, generating control commands for multiple actuators may include generating multiple control commands and selecting one or more control commands that are closest to at least one reference command.

[0133] In some embodiments, system 1000 may further generate control commands for multiple actuators based on one or more temperatures associated with one or more aircraft components (e.g., engines, battery packs). For example, system 1000 may further generate control commands for multiple actuators based on one or more temperatures associated with one or more engines, such that one or more temperatures are maintained within one or more predetermined ranges (e.g., continuous range, takeoff / landing range, emergency range). For example, system 1000 may generate control commands that adjust weights or other parameters to reduce the torque commanded to one or more engines. In some embodiments, system 1000 may generate control commands for multiple actuators such that one or more temperatures associated with one or more aircraft components do not exceed predetermined temperatures (e.g., operational limits, transient failure limits, warning limits).

[0134] In some embodiments, system 1000 may generate control commands for multiple actuators based on priorities associated with energy optimization and temperature management. 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 management over energy optimization. Alternatively, based on the second remaining time being greater than the first remaining time, system 1000 may prioritize energy optimization over temperature management.

[0135] In step 1212, system 1000 can control multiple actuators to meet (e.g., respond to, comply with, resolve, or be based on) desired commands of the electric aircraft based on the generated control commands. For example, system 1000 can send at least one (e.g., each) of the generated control commands to its corresponding actuator, causing the actuator to move according to the control commands.

[0136] In some embodiments, method 1200 can result in optimized energy use, maximizing the amount of energy available on the aircraft. In some embodiments, method 1200 may cause all battery packs to deplete their energy simultaneously. In fact, method 1200 may maximize the available flight range and can also increase safety by preventing dangerous situations where one or more batteries deplete their energy before the others and are no longer able to power their respective EPUs.

[0137] According to some embodiments, Figure 13A An exemplary battery pack failure scenario without energy optimization is shown, and Figure 13B An exemplary battery pack failure scenario with energy optimization (e.g., using system 1100 and / or method 1200) is illustrated. In this exemplary scenario, both figures show the bus energy and power consumption of the aircraft's three buses (each bus connected to two batteries) at different points in the flight or trip (e.g., the start point, midway through the flight distance, and the end point), where, at the midway point, the battery pack connected to bus 1 fails. Figure 13A As shown, because there is no energy optimization, even after a battery failure on bus 1, the same amount of power is drawn from each bus (e.g., among buses 1, 2, and 3), causing bus 1 to run out of power at approximately 3 / 4 of the total distance traveled. Conversely, Figure 13B The diagram illustrates energy optimization, where, after a battery pack connected to bus 1 fails, the energy optimization function adjusts the amount of power drawn from each bus to cope with the battery pack failure. This allows the aircraft to reach approximately 11 / 12 of the mission distance and simultaneously deplete all battery packs. Energy optimization is achieved, for example, by using... Figure 11 or Figure 12 The technology described allows aircraft to extend their flight range by allocating energy or power drawn from battery packs.

[0138] According to some embodiments, Figure 13C An exemplary battery pack failure scenario without energy optimization is shown, and Figure 13DAn exemplary battery pack failure scenario with energy optimization (e.g., using system 1100 and / or method 1200) is illustrated. In this exemplary scenario, both figures show the bus energy and power consumption of the aircraft's three buses (each bus connected to two batteries) at different flight distances, where, at a midway point, the battery pack connected to one of the buses fails. Figure 13C As shown, if energy optimization is not implemented and the system starts with 50% buffered energy (e.g., 50% of nominal energy) in addition to the nominal energy, then even after a battery failure on bus 1, the same amount of power is drawn from each bus, causing one bus to run out of power, and both buses still having 50% remaining energy at the end of the task. Conversely, Figure 13D Energy optimization was implemented, starting with a 10% buffer energy in addition to the nominal energy, and the amount of power drawn from each bus was adjusted to handle battery pack failures. This allowed the aircraft to reach the end of its journey with less buffer energy used than without energy optimization. Figure 11 or Figure 12 The described technology can complete the task with less buffer energy.

[0139] In some embodiments, the energy optimization function can be configured to receive battery information (e.g., state of charge, state of energy, state of power, state of health, state of temperature, usable energy, etc.) from the battery management system (BMS) of the flight control system for at least one (e.g., each) battery. The state of charge (SOC) can refer to the available battery pack capacity relative to the rated capacity of the battery pack. For example, the state of charge can be based on the open-circuit voltage (OCV) of the battery pack, where OCV is the resting voltage of the battery pack (e.g., based on a battery pack with no current flowing over a set period of time). The state of energy (SOE) can be a measure of the remaining discharged energy of the battery. The state of power (SOP) can be the maximum power capacity that the battery can deliver over a period of time. The state of health (SOH) can refer to the overall condition of the battery compared to its ideal condition (e.g., the amount of degradation). The state of temperature (SOT) can refer to the current temperature of the battery. Based on the received battery information for at least one (e.g., each) battery and the high-voltage architecture of the flight control system, the energy optimization function can set or adjust a reference point (e.g., a reference command) for the flight control system to perform control assignments.

[0140] Figures 14 to 18Exemplary scenarios illustrating the implementation of energy optimization functions (e.g., all or part of system 1100 and / or method 1200) according to some embodiments are shown. In each scenario, HV bus 1 is associated with batteries 1 and 4 and 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 and 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 and 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.

[0141] "Battery index" (e.g., Figures 14 to 18 The battery indices 1402, 1502, 1602, 1702, and 1802 can indicate the available energy associated with each battery pack (e.g., batt1, batt2, batt3, batt4, batt5, batt6). The "HV bus energy index" (e.g., ...) indicates the available energy associated with each battery pack. Figures 14 to 18 The HV bus power indices (1404, 1504, 1604, 1704, 1804) can be shown for each HV bus (e.g., bus1, bus2, bus3) and / or HV channel (e.g., ...). Figure 18 Available energy associated with ch1 and ch4. "Engine index" (e.g., Figures 14 to 18 The engine indices 1406, 1506, 1606, 1706, and 1806 can indicate 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., Figures 14 to 18 The HV bus power indices (1408, 1508, 1608, 1708, 1808) can be shown for use with each HV bus (e.g., bus1, bus2, bus3) and / or HV channel (e.g., ...). Figure 18 The initial reference command (e.g., "input reference") and the adjusted reference command (e.g., "output reference") associated with power consumption in ch1 and ch4.

[0142] Figure 14The nominal scenario 1400 is shown, in which each battery pack of batteries 1-6 has equivalent available energy (e.g., indicated by 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 to the energy optimization function (e.g., "input reference") can be the same as the reference command output / adjusted by the energy optimization function (e.g., "output reference"), as shown by engine index 1406 and HV bus power index 1408.

[0143] Figure 15 The diagram illustrates a battery failure scenario 1500, where battery batt1 has failed, and all other batteries (e.g., batt2-batt6) have equivalent available energy (indicated by battery index 1502). The HV bus bus1 associated with the failed batt1 has approximately half the available energy of each of HV buses bus2 and bus3 (indicated by HV bus energy index 1504). As shown in engine index 1506, the energy optimization function adjusts the reference command (e.g., "output reference") associated with engines EE1, EE4, EE9, and EE12 (powered only by battery batt2 via HV bus bus1) to a lower value (e.g., "output reference" approximately 0.6), and adjusts the reference command associated with engines EE2, EE3, EE5-EE8, EE10, and EE11 to a higher value (e.g., "output reference" approximately 1.2). As shown in HV bus power index 1508, the energy optimization function can adjust the reference command associated with HV bus bus1 connected to the faulty battery pack batt1 (e.g., "Output Reference" is 2.4) to a lower value, and increase each reference command associated with HV buses bus2 and bus3 (e.g., "Output Reference" is 4.8 for both buses 2 and 3) to a higher value, in order to reduce power consumption from bus bus1 and increase power consumption from buses bus2 and bus3. This prevents HV bus bus1 from running out of power before HV buses bus2 and bus3.

[0144] Figure 16The diagram illustrates engine failure scenario 1600, where 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 shown in engine index 1606, the output reference power command associated with engine EE1 (e.g., "output reference") is reduced to 0, and the energy optimization function adjusts the reference power commands associated with engines EE4, EE9, and EE12 to a higher level (e.g., the "output reference" associated with 4, 9, and 12) to increase the torque / speed / usage associated with engines EE4, EE9, and EE12. The energy optimization function adjusts the reference power commands associated with engines EE2, EE3, EE5-EE8, EE10, and EE11 to a lower level (e.g., the “output reference” associated with 2, 3, 5-8, 10, and 11) to reduce the torque / speed / usage associated with engines EE2, EE3, EE5-EE8, EE10, and EE11. As shown in HV bus power index 1608, the energy optimization function can adjust the reference commands so that even if HV bus bus1 connected to the failed engine EE1 is only used to power engine EE12, the output reference commands associated with each bus are approximately the same (e.g., the “output reference” for buses 1-3). As shown in scenario 1600, in response to an engine failure, the energy optimization function can adjust the reference commands associated with the engine so that the remaining energy or remaining discharge time of each HV bus is approximately the same (e.g., so that the difference is within a threshold).

[0145] Figure 17 The diagram illustrates HV channel failure scenario 1700, where HV channel 1 of bus 1 has failed. Due to the HV channel 1 failure, the available energy of battery batt1 may drop to 0 (indicated by battery index 1702), and HV bus bus 1 may have only half the available energy of each of buses bus 2 or bus 3 (indicated by HV bus energy index 1704). As shown in engine index 1706, the energy optimization function may not need to adjust the engine reference command because the HV channel 1 failure causes the reference commands associated with engines EE1 and EE12 to shrink to 0, which may cause the reference commands associated with bus bus 1 to decrease automatically (as shown in HV bus power index 1708). As shown in scenario 1700, the impact of losing one battery can be mitigated (e.g., offset) by the impact of losing two engines.

[0146] Figure 18The diagram illustrates a crosslink failure scenario 1800, where a crosslink (e.g., HV bus bus1) connecting battery packs batt1 and batt4 has failed. Due to the crosslink failure, battery batt1, having only half its available energy (indicated by battery index 1802), may cause HV channel 1 ch1 (of the failed crosslink) to have only one-quarter of the available energy of each of buses bus2 or bus3. Additionally, due to the crosslink failure, battery batt4, having the same available energy (indicated by battery index 1802) as batteries batt2, batt3, batt5, and batt6, may cause HV channel 4 ch4 (of the failed crosslink) to have only half the available energy of each of buses bus2 or bus3 (indicated by HV bus index 1804). As shown in engine index 1806, the energy optimization function can adjust the reference command by using the estimated remaining discharge time per bus / channel. For example, the energy optimization function adjusts the reference commands associated with engines EE1 and EE12 (which are powered solely by battery batt1 via HV channel ch1) to a lower level (e.g., "output reference" approximately 0.6) to reduce the torque / speed / usage associated with engines EE1 and EE12, and adjusts the reference commands associated with engines EE2-EE11 to a higher level (e.g., "output reference" approximately 1.2) to increase the torque / speed / usage associated with engines EE2-EE11. As shown in HV channel index 1808, the energy optimization function can reduce the reference commands associated with HV channel ch1 (the faulty crosslinker) connected to battery batt1, and can increase each reference command 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.

[0147] In some embodiments, the flight control system may include one or more engine thermal management functions. Engine thermal imbalance may have several causes, such as uneven propeller command, differences in heat dissipation, differences in battery and / or engine health, and battery or engine failure. Electric engines may have thermal operating limits, and engine thermal management functions may be configured to prevent the engine from experiencing any unacceptable operating characteristics (e.g., operation exceeding risk thresholds) and to prevent the engine from exceeding operating limits, including in-failure conditions.

[0148] In some embodiments, at least one (e.g., each) electric motor (e.g., at least one processor associated with each motor (e.g., a portion of each motor and / or connected to each motor)) may be configured to estimate the remaining time of the motor at the current power setting (e.g., as part of system 1100 and / or method 1200). In some embodiments, at least one electric motor may include an engine time estimator (e.g., using an inverter or processor of the electric motor) 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 (e.g., control board, DC link capacitor, fast discharge, motor control unit (MCU), power module connector, stator winding, oil, or rotor magnet of EPU). In some embodiments, the engine time estimator may normalize the temperature in terms of time. In some embodiments, the engine time estimator may use time-based temperature metrics to normalize the determined (e.g., calculated, measured, or received from) temperatures for different components (e.g., rotor magnet, oil, control board, DC link capacitor, motor control unit (MCU), fast discharge, power module connector, stator winding, etc.) with different time constants and temperature limits. For example, different components of an electric motor 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 with the least estimated remaining time. For example, the engine time estimator may transmit the lowest estimated remaining time as the estimated remaining time of the engine at the current power setting to the FCS.

[0149] Figure 19 Exemplary temperature-to-time graphs are shown for components used to estimate remaining time (e.g., determined and / or used as part of system 1100 and / or method 1200) according to various embodiments. An engine time estimator can generate graphs for one or more components to aggregate the temperatures of different components with different temperature limits into a time-based metric, thereby enabling visualization and / or data structures (e.g., usable by engine thermal management functions) of the engine's minimum overheating time. 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 flight phase may have a different power output.

[0150] Figure 20 Exemplary meters are shown, according to various embodiments, for displaying information related to the estimated remaining time (e.g., determined and / or used as part of system 1100 and / or method 1200). In some embodiments, the meters may display the proximity of torque or other aircraft parameters to defined limits. Although Figure 20 The meters are shown as a combination of certain layout and visual elements; however, it should be understood that other variations can be used to indicate the proximity of the current operating aircraft state to the limits and / or the amount of time remaining in the operating aircraft state (e.g., safe time, remaining time until the engines reach or exceed limits). In some embodiments, the information displayed in the meters may be based on information regarding... Figure 19 or Figures 21 to 2 Information derived from the features described in one or more of the figures in Figure 5. Meters (e.g., meters 2002, 2004, 2006) may include a torque arc (e.g., torque arc 2008) showing the available torque from the running engine, an indicator showing the torque in use (e.g., indicator 2010), a limit mark showing the current effective torque limit (e.g., continuous limit mark 2012 (CONT), takeoff limit mark 2014 (TO), emergency limit mark 2016 (EMRG)), a red line showing the maximum torque (e.g., red line 2018), and an enhancement bar showing the remaining time at the current rated maximum torque (e.g., enhancement bar 2020). For example, meter 2002 displays continuous limit mark 2012 to indicate that the current effective torque limit is a continuous limit associated with an infinite remaining time as shown by enhancement bar 2022. Meter 2004 displays takeoff limit marker 2014 to indicate that the current effective torque limit is a takeoff limit associated with the remaining time below a first predetermined threshold (e.g., 243 seconds, below 300 seconds), as indicated by reinforcement bar 2020. Meter 2006 displays emergency limit marker 2016 to indicate that the current effective torque limit is an emergency limit associated with the remaining time below a second predetermined threshold (e.g., 27 seconds, below 30 seconds), as indicated by reinforcement bar 2024. In some embodiments, the second predetermined threshold includes the time less than the first predetermined threshold.

[0151] In some embodiments, the engine time estimator can be configured to determine one or more temperature parameters. For example, the engine time estimator can 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 can 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 can be configured to verify one or more communication signals (e.g., checking for faults, loss of communication within a timeout window, etc.). In some embodiments, the engine time estimator can be configured to calculate a composite temperature value based on the validity of the inputs.

[0152] In some embodiments, the engine timing estimator may be configured to detect one or more temperature faults (e.g., when the temperature reaches the FCS warning threshold, when the temperature reaches the engine operating level, thermal faults on the power stage, low oil level, low oil flow, no inverter oil flow, no stator oil flow, no gearbox oil flow, HEX airflow blockage, temperature sensor degradation, temperature sensing loss, single-phase short circuit, etc.). For example, the engine timing estimator may be configured to perform a validity check on the composite temperature value. In some embodiments, detecting one or more temperature faults may include determining whether the composite temperature value is outside the rated range. In some embodiments, detecting one or more temperature faults may include determining whether the composite temperature value is outside the limit range. In some embodiments, detecting one or more temperature faults may include determining whether the engine has thermally deteriorated (e.g., excessive increase in ΔT due to oil leaks).

[0153] 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 magnet, oil, control board, DC link capacitor, motor control unit (MCU), fast discharge, power module connector, stator winding, etc.). In some embodiments, the background thermal model may utilize multiple nodes to predict the temperature of at least one component of the engine. Alternatively, the background thermal model may utilize a detailed loss function to predict the temperature of at least one critical component of the engine. In some embodiments, the background thermal model may be fine-tuned and / or undergo multiple rounds of testing to improve the accuracy of the predicted temperature.

[0154] In some embodiments, the engine time estimator may run a remaining time model. For example, the remaining time 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 remaining time model may perform forward prediction to determine (e.g., predict) the remaining time until one or more thermal limits of at least one (e.g., each) component of the engine are reached, based on the determined current thermal state of the component. In some embodiments, the remaining time model may take into account external environmental conditions. In some embodiments, the remaining time model may include a reduced-order thermal model with a simplified loss function. In some embodiments, the remaining time model may include a flattened analytical fit exponentially rising curve for at least one (e.g., each) engine component of interest (e.g., rotor magnet, oil, control board, DC link capacitor, motor control unit (MCU), fast discharge, power module connector, stator winding, etc.). In some embodiments, the remaining time model may include a machine learning model (e.g., a neural network model and / or a model trained with engine temperature values, flight parameters, and / or energy states of one or more flight phases), said machine learning model being trained to take 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 can generate a machine learning model based on training data (e.g., simulation, historical, data from a background thermal model).

[0155] In some embodiments, at least one of the engine thermal management function or flight control system may be configured to perform one or more preventative actions to prevent overheating. In some embodiments, the preventative actions may include one or more of the following: transmitting a warning (e.g., to the flight control system, to the engine, to the aircraft pilot, or via GPS to the nearest landing area), performing a torque descent (e.g., reducing the torque required for at least one EPU over time and / or adjusting at least one command, etc.), or performing an emergency shutdown.

[0156] In some embodiments, at least one (e.g., each, multiple) engine can transmit its corresponding estimated remaining time to the flight control system.

[0157] Figure 21Exemplary temperature thresholds are shown according to various embodiments (e.g., applicable to system 2200, method 2300, and / or method 2400). Threshold 2131 may be a maximum rated point / soft upper limit, whereby the engine can send an alert to the FCC when engine parameters exceed threshold 2131. In response to receiving the alert, the FCC may perform one or more preventative actions. For example, the FCC may output an alert to the pilot, command the engine to reduce torque output, and / or command the engine to shut down. Exceeding threshold 2132 may indicate that the engine is approaching an operational limit, and the flight control system may automatically reduce the engine's torque output in response to exceeding this threshold to prevent engine overheating. Threshold 2133 may be an operational limit, and the flight control system may perform 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 two pilot controls, buttons, switches) may be configured to exceed the operational limit threshold and / or trigger one or more emergency actions performed by the flight control system.

[0158] Figure 22 This is a functional block diagram of an exemplary system 2200 for a VTOL aircraft that includes torque limit calculation functionality (e.g., torque limit calculation function 2226), consistent with the disclosed embodiments. It should be understood that specific machines (e.g., aircraft) can use the exemplary system 2200 to implement improvements in the technical fields of aircraft safety, stability, reliability, and efficiency. For example, consistent with the disclosed embodiments, some embodiments may involve dynamically changing torque commands based on at least one aircraft condition, which can in turn be used by the propulsion unit to help make the aircraft, for example, safer, more stable, easier to fly, more reliable, and more efficient during different phases or modes of flight. Figure 22 As shown, system 2200 may include torque limit calculation function 2210, control distribution function 2220, and DCPS 2230. System 2200 may be implemented by a microprocessor-based controller storing in a storage medium executing software code to implement the functions described herein. System 2200 may also be implemented in hardware or a combination of hardware and software. System 2200 may be configured to repeatedly execute single steps or sequences until a desired or commanded result is achieved. It should be understood that, for ease of description, Figure 22 Many of the conventional functions of the control system are not shown in the diagram.

[0159] In some implementations, the torque limit calculation function 2210 may be configured to receive at least one of system malfunction, pilot input, or engine data as input. For example, the torque limit calculation function 2210 may be configured to receive (e.g., from multiple engines or each engine of the aircraft) one or more of the following: a predetermined propeller torque limit 2201, an engine mode 2202, an engine malfunction state 2203, an engine continuous torque limit 2204, an engine takeoff torque limit 2205, an engine emergency torque limit 2206, an engine overheat torque limit 2207, an estimated range at maximum takeoff torque 2208, or an estimated range at maximum emergency torque 2209. In some embodiments, one or more inputs may be preset based on at least one of the aircraft's design specifications, structural limitations, or cooling capabilities. For example, one or more inputs may be associated with regulatory values.

[0160] In some embodiments, system 2200 may include engine rated power management function 2215. Engine rated power management function 2215 may be configured to select an appropriate torque rating (e.g., engine continuous torque limit 2204, engine takeoff torque limit 2205, engine emergency torque limit 2206) for at least one engine (e.g., each or more engines of an aircraft) based on flight phase (e.g., hovering, cruise / continuous, transition, takeoff / landing) and one or more operational requirements. For example, function 2215 may be configured to select engine takeoff torque limit 2205 as the torque rating based on determining that the aircraft is in the takeoff or landing phase of flight. Alternatively, function 2215 may be configured to select engine continuous torque limit 2204 as the torque rating based on determining that the aircraft is in the cruise phase of flight. Alternatively, function 2215 may be configured to select engine emergency torque limit 2206 as the torque rating based on detecting one or more system faults (e.g., engine failure, damaged aircraft components, etc.). In some embodiments, the management rated power function 2215 may include an automatic function configured to dynamically switch between torque ratings based on flight phase, operational requirements, and aircraft dynamics (e.g., aircraft condition, system failure). In some embodiments, the management engine rated power function 2215 may include a switch configured to receive pilot input via a switching feature.

[0161] In some embodiments, system 2200 may include engine thermal management function 2221. Engine thermal management function 2221 may be configured to manage (e.g., adjust, configure, set, and / or monitor) engine temperature to ensure (e.g., each, at least one) engine operates within safe parameters, thereby preventing engine components from overheating and being damaged. For example, engine thermal management function 2221 may be configured to determine engine rated torque limit 2216 for at least one aircraft engine based on the maximum power output from engine rated power management function 2215 (e.g., for each engine or multiple aircraft engines).

[0162] Engine rated torque limit 2216 may correspond to the maximum torque that the engine can produce under normal continuous operating conditions. In some embodiments, engine thermal management function 2221 may receive one or more of the following from at least one (e.g., each) engine: engine overheat torque limit 2207, expected range at maximum takeoff torque 2208, or expected range at maximum emergency torque 2209. Based on at least one of the expected range at maximum takeoff torque 2208 or expected range at maximum emergency torque 2209, engine thermal management function 2221 may determine which limit (e.g., engine continuous torque limit 2204, engine takeoff torque limit 2205, engine emergency torque limit 2206, engine overheat torque limit 2207, engine rated torque limit 2216) to determine steady-state engine torque limit 2222. For example, the determined amount of remaining time for takeoff torque can be represented by a timer indicating the remaining time at 2217 (e.g., greater than 0, indicating the aircraft can maintain its operation at maximum takeoff torque) and the expected endurance 2208 at maximum takeoff torque, with the remaining time at 2218. Based on the determined amount of remaining time for takeoff torque, engine thermal management function 2221 can use engine rated torque limit 2216 to determine steady-state engine torque limit 2222. Alternatively, the determined amount of no remaining time for takeoff torque can be represented by a timer indicating the no remaining time at 2217 (e.g., not greater than 0) and the expected endurance 2208 at maximum takeoff torque, with the remaining time at 2218. Based on the determined amount of no remaining time for takeoff torque, engine thermal management function 2221 can use engine continuous torque limit 2204 to determine steady-state engine torque limit 2222. Alternatively, the timeout of emergency torque can be represented by a timer 2209 indicating the expected endurance at maximum emergency torque (e.g., not greater than 0, at which point the aircraft cannot maintain its operation at maximum emergency torque) at 2218. Based on the timeout of emergency torque, engine thermal management function 2221 can use engine overheat torque limit 2207 to determine steady-state engine torque limit 2222. In some embodiments, engine thermal management function 2221 can be configured such that there is a time buffer between timer 2217 and timer 2218 (e.g., so that timer 2217 and timer 2218 do not reach 0 simultaneously).

[0163] In some embodiments, engine thermal management function 2221 may be configured to return the torque limit to its normal torque rating upon receiving an indication of engine cooling (e.g., based on receiving an indication of engine cooling, in response to receiving an indication of engine cooling). For example, after setting engine overheat torque limit 2207 as the engine's torque limit, engine thermal management function 2221 may receive an indication that the engine temperature has decreased below a predetermined temperature threshold. Alternatively, engine thermal management function 2221 may receive an indication that the remaining time of the expected range at the torque rating (e.g., timer 2218) is greater than a predetermined time threshold (e.g., indicating sufficient engine cooling). Engine thermal management function 2221 may be configured to set one of torque limits 2204, 2205, or 2206 as the engine's torque limit based on said indication.

[0164] 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 predetermined propeller torque limits. In some embodiments, one or more predetermined 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, structural load limits may be associated with loads applied by one or more of the following: the engine on the propeller hub, the rotation of the propeller, the propeller blades, the support between the propeller and the engine, the support between the engine and the boom, the support between the boom and the wing, or the boom itself. Alternatively or additionally, one or more predetermined propeller torque limits 2201 may include one or more predetermined values ​​specified for a particular aircraft configuration. For example, one or more predetermined 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, 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 engine modes 2202 (e.g., the state of the inverter's internal state machine, factor 0 in standby / listen mode, factor 0 in retracted mode, factor 1 in closed-loop torque command mode, etc.) or engine fault states 2203. In some embodiments, engine availability determination 2211 may output engine availability factor 2212, which may indicate the level of functionality associated with one or more engines (e.g., regarding expected or full functionality). For example, engine availability factor 2212 may include a value such as 0 for inoperability, 1 for full operation, and 0.5 for a fault condition (e.g., the loss of one of two sets of windings resulting in only about half the normal torque output of the engine). In some embodiments, based on the torque limits determined by the engine thermal management function 2221 (e.g., engine continuous torque limit 2204, engine rated torque limit 2216, engine overheat torque limit 2207), the torque limit calculation function 2210 can combine (e.g., multiply) the engine availability factor 2212 with the torque limits at 2213. At 2214, the torque limit calculation function 2210 can compare the outputs of the predetermined propeller torque limit 2201 and 2213, and can output the lower of the two as the steady-state engine torque limit 2222.

[0165] 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 combining (e.g., multiplying) the engine availability factor 2212 output from engine availability determination 2211 with the engine emergency torque limit 2206 at 2219. In some embodiments, the steady-state engine torque limit 2222 for the engine may include a lower limit than the dynamic engine torque limit 2224 for the engine.

[0166] The control allocation function 2220 can 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 can be configured to determine one or more allocated torque commands 2223 based on at least one of a steady-state engine torque limit 2222, an estimated endurance 2208 at maximum takeoff torque, or an estimated endurance 2209 at maximum emergency torque.

[0167] DCPS 2230 can 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, DCPS 2230 can 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.

[0168] Figure 23 This is a flowchart illustrating an exemplary method 2300 for engine thermal management of a specific machine (e.g., an aircraft) according to some embodiments of the present disclosure, thereby improving the aircraft's technology (e.g., aircraft safety, thermal efficiency, range capability, payload capacity, structural integrity). The steps of method 2300 may be performed by, for example, in... Figures 1 to 8 , Figures 9A to 9E , Figures 10 to 12 , Figures 13A to 13D , Figures 14 to 22 , Figure 24 Performed or otherwise used on any component of the aircraft shown in Figure 25. Figure 10 The method may be executed by system 1000, or by any flight control computer (e.g., a computer-implemented method) or flight control system. For example, the aircraft's flight control computer may be configured to execute one or more steps of method 2300.

[0169] The steps of method 2300 can be activated or adjusted as needed based on at least one aircraft flight condition, at least one maneuver, and / or at least one operational requirement. It should be understood that the illustrated method 2300 can be modified to alter the order of steps and include additional steps. It should also be understood that the complexity of method 2300 would make it impossible, or at least highly impractical, for a human user to implement the method effectively, especially when considering the implementation of these functions while the aircraft is in flight (including takeoff or landing) (e.g., in real time), and the energy used by the aircraft and the constantly changing conditions of the batteries and engines. Furthermore, engine thermal management (e.g., using method 2300) improves the aircraft's safety and performance capabilities (e.g., the extent to which the engine can approach maximum performance without exceeding thermal limits) and the lifespan of different components (e.g., batteries, engines) without any direct intervention from the user (e.g., the pilot). In some embodiments, the aircraft of system 1000 may include an electric aircraft comprising multiple actuators, each actuator comprising multiple electric propulsion units (EPUs) and multiple battery packs powering the multiple electric propulsion units.

[0170] In step 2302, system 1000 may determine one or more desired commands for an electric aircraft. For example, the one or more desired commands may include at least one of the following: one or more force or torque commands (e.g., Figure 10 The inputs to the control assignment 1029), desired position and / or turning rate commands (e.g., Figure 10 The output of the turning rate command model 1004), the desired position and / or lateral speed command (e.g., Figure 10 The output of the lateral velocity command model 1006), one or more of the desired height, vertical velocity, or vertical acceleration commands (e.g., Figure 10 The output of the climb command model 1008), the desired position, longitudinal velocity, or longitudinal acceleration command (e.g., Figure 10 The output of the forward speed command model 1010), and alternative commands used to achieve the desired change (e.g., based on such as...). Figure 10 The feedback discussed in the article), one or more total expectancy (e.g., based on Figure 10 The outputs of feedback 1012, 1016, 1018, 1022 and feedforward 1014, 1020 are calculated), and one or more torque commands (e.g., by...) Figure 10The inner loop control law 1028 is determined), linear command, angular acceleration command, or any command in the form of a signal (e.g., pilot input from a pilot input device, input from an autopilot system, a signal generated by the flight control system or a computer) associated with the desired motion of the aircraft (e.g., the motion desired by the pilot of the aircraft, the motion desired by the flight control system, the motion desired by the autopilot system).

[0171] In step 2304, system 1000 may receive engine information from at least one (e.g., each) of a plurality of EPUs. For example, system 1000 may also monitor the engine status of at least one (e.g., each) of the plurality of electric propulsion units. For example, system 1000 may receive engine information (e.g., engine status) associated with at least one (e.g., each) of the plurality of EPUs (e.g., via a digital communication interface, via an inverter and / or processor of the EPU). The 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 relative to a limit), one or more limits associated with at least one engine (e.g., temperature limit), the time associated with the temperature of at least one (e.g., each) engine (e.g., remaining time at current ratings such as continuous, takeoff, or emergency, remaining time until a limit (e.g., thermal limit) is reached), or a status (e.g., fault condition, engine is currently running (e.g., activated) and available for control, engine is currently not running and unavailable for control). In some embodiments, engine information may include one or more of the following: engine overheat torque limit, expected range at maximum takeoff torque, or expected range at maximum emergency torque.

[0172] In step 2306, system 1000 may generate control commands for multiple actuators based on the received engine information. For example, 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, thereby preventing engine components from overheating and being damaged. In some embodiments, system 1000 may generate control commands that limit operations performed by at least one engine or other component of the aircraft to reduce strain or risk to associated components while still controlling the actuators based on desired commands (e.g., minimizing the impact on reference states, flight envelope, etc.). In some embodiments, system 1000 may determine engine rated torque limits for at least one (e.g., each) engine. Engine rated limits may include the maximum torque that the engine is capable of generating under normal (e.g., continuous) operating conditions. In some embodiments, system 1000 may determine engine rated torque limits based on the received engine information. Figure 22 The description provides a detailed description of generating control commands (e.g., assigned torque command 2223 or torque command 2232). In some embodiments, generating control commands for multiple actuators may further be based on one or more aircraft conditions. For example, system 1000 may generate control commands based on the current aircraft condition.

[0173] In some embodiments, the generation of control commands may be further based on the state of energy of multiple battery packs. For example, system 1000 may monitor the state of energy 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., physical connection, such as a bus and / or channel system, or via a digital communication interface). In some embodiments, the aircraft may contain a single BMS configured to manage all battery packs on the aircraft. In some embodiments, at least one (e.g., each) battery pack may contain its own BMS. In some embodiments, system 1000 may receive battery information (e.g., state of energy) associated with at least one (e.g., each) battery from one or more BMSs associated with multiple battery packs. Battery information may include one or more of the following: available energy of at least one (e.g., each) battery pack, remaining discharge time (e.g., at least one (e.g., each) crosslinker, at least one (e.g., each) independent high-voltage bus, at least one (e.g., each) high-voltage channel), state of energy (SOE), state of charge (SOC), state of power (SOP), state of health (SOH), fault condition (e.g., short circuit or overcurrent, whether the battery is activated / functioning), or state of temperature (SOT).

[0174] In step 2308, system 1000 can control multiple actuators to meet (e.g., respond to, comply with, resolve, or be based on) desired commands of the electric aircraft based on the generated control commands. For example, system 1000 can send at least one (e.g., each) of the generated control commands to its corresponding actuator, which can move based on (e.g., according to) the control commands.

[0175] In some embodiments, method 2300 can produce optimized engine performance by maximizing the performance of at least one (e.g., each) engine while remaining within engine thermal limits. In effect, method 2300 can maximize availability and also increase safety by preventing dangerous situations where one or more engines shut down before others due to thermal overload and are no longer able to generate thrust.

[0176] Figure 24 This is a block diagram illustrating an exemplary method 2400 for engine thermal management and energy optimization for a particular machine (e.g., an aircraft) according to various embodiments of the present disclosure, thereby improving the aircraft's technology (e.g., aircraft safety, thermal efficiency, range capability, payload capacity, structural integrity). The steps of method 2400 can be, for example, in... Figures 1 to 8 , Figures 9A to 9E , Figures 10 to 12 , Figures 13A to 13D , Figures 14 to 23 Performed or otherwise used on any component of the aircraft shown in Figure 25. Figure 10 The method can be executed by system 1000, or by any flight control computer (e.g., a computer-implemented method) or flight control system. For example, the aircraft's flight control computer can be configured to execute one or more steps of method 2400. It should also be understood that the complexity of method 2400 would make it impossible, or at least highly impractical, for a human user to implement the method effectively, especially when considering the implementation of these functions while the aircraft is in flight (including takeoff or landing) (e.g., in real time), and when the energy used by the aircraft and the battery status are constantly changing. Furthermore, energy optimization (e.g., using method 1200) improves the aircraft's safety and range capabilities (e.g., how much energy is available to power the aircraft) and the lifespan of different components (e.g., batteries) without requiring any direct involvement from the user (e.g., the pilot). It should be further understood that, although... Figure 24 Examples of energy optimization and thermal management related to torque and power are depicted, but Figure 24This may be alternatively or additionally associated with any parameters (such as speed and / or current) related to the aircraft's actuators. The steps of method 2400 may be activated or adjusted as needed based on the aircraft's flight conditions, maneuvers, or operational requirements. It should be understood that method 2400 may be altered to modify the order of steps and include additional steps.

[0177] In step 2401, the thermal management function may calculate at least one torque or power adjustment factor 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 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) engine. In some embodiments, the thermal management function may calculate at least one torque or power adjustment factor based on one or more of the received battery information, EWIS information, or engine information.

[0178] In step 2402, the energy optimization function can calculate a power adjustment factor using, for example, the difference in available energy (AE) per bus. In some embodiments, the energy optimization function can receive at least one crosslink state from a plurality of crosslink states. In some embodiments, the energy optimization function can receive the difference in available energy per bus from at least one BMS.

[0179] In step 2403, the flight control system can be configured to adjust at least one electric engine reference value (e.g., a reference command) based on values ​​calculated by the thermal management function and the energy optimization function. In some embodiments, the flight control system can determine which control assignment should be solved for a new reference state. In some embodiments, the flight control system can calculate the deviation (e.g., difference) between the current state (e.g., configuration, actuator position, actuator output) and the new reference state (e.g., associated with at least one electric engine reference value, reference torque setting).

[0180] In 2404, the flight control system can be configured to recalculate force / torque commands. (The above...) Figure 11 The description provides a detailed description of the recalculate force / torque command.

[0181] Figure 25A and Figure 25B Exemplary scenarios are shown where both thermal management and energy optimization are performed according to some embodiments (e.g., using and / or according to system 1100, method 1200, method 2300, and / or method 2400). Figure 25A As shown in Figure 2500a, all electric engines start with the same initial torque reference value (e.g., reference command). Upon startup, HV bus 2 has more energy than bus 1 and bus 3 (indicated by B2 and B5 being more fully charged than B1 and B4, and B3 and B6), and electric motor (EE) 8 is hotter than the other motors. Figure 25B As shown in Figure 2500b, thermal management and energy optimization are performed by adjusting the initial torque reference values. For example, the reference values ​​associated with electric engines 2, 5, and 11 on bus 2, which have the most available energy, are increased (e.g., to increase the use of electric engines 2, 5, and 11), while the reference value associated with electric engine 8 is decreased (e.g., to reduce its use relative to other engines). Additionally, the reference values ​​associated with the electric engines on buses 1 and 3 are adjusted such that the utilization rate of the engine associated with each of buses 1 and 3 is lower than that of engines 2, 5, and 11. Performing both thermal management and energy optimization results in full energy utilization across all HV buses while remaining within engine thermal limits.

[0182] In some embodiments, engine thermal management may take precedence over energy optimization. Alternatively, energy optimization may take precedence over engine thermal management.

[0183] Additional aspects of this disclosure may be further described by the following terms: 1. A computer-implemented method, comprising: Determine one or more desired commands for the aircraft; At least one reference command is determined based on the one or more desired commands and one or more aircraft conditions; Monitor the energy status of multiple battery packs of the aircraft, wherein at least a first battery pack of the multiple battery packs is electrically isolated from at least a second battery pack of the multiple battery packs; The at least one reference command is adjusted based on the monitored state of energy of the plurality of battery packs; Based on at least one adjusted reference command, control commands for multiple actuators of the aircraft are generated; and The multiple actuators are controlled according to the generated control commands to satisfy one or more desired commands of the aircraft.

[0184] 2. The computer-implemented method according to Clause 1, wherein the monitored state of energy includes the usable energy of the plurality of battery packs.

[0185] 3. The computer-implemented method according to Clause 1 or 2, wherein the monitored energy state includes the remaining discharge time associated with one or more high-voltage channels.

[0186] 4. A computer-implemented method according to any one of clauses 1 to 3, wherein the monitored state of energy includes the difference in state of energy between at least two of the plurality of battery packs.

[0187] 5. A computer-implemented method according to any one of Clauses 1 to 4, wherein the monitored energy state includes the difference in power consumption between at least the first engine and the second engine of the plurality of electric propulsion units of the aircraft.

[0188] 6. The computer-implemented method according to Clause 5, wherein the first engine has lower power consumption than the second engine, and wherein the generated control command causes a reduction in the power consumption of the second engine.

[0189] 7. A computer-implemented method according to any one of Clauses 1 to 6, wherein the adjustment of the at least one reference command is further based on the state of one or more engines of a plurality of electric propulsion units of the aircraft.

[0190] 8. A computer-implemented method according to any one of Clauses 1 to 7, wherein adjusting the at least one reference command comprises updating one or more reference commands to optimize the remaining discharge time across one or more high-voltage channels.

[0191] 9. A computer-implemented method according to any one of Clauses 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.

[0192] 10. A computer-implemented method according to any one of Clauses 1 to 9, wherein the one or more aircraft conditions include one or more of aircraft dynamics, flight conditions, or the condition of at least one aircraft component.

[0193] 11. A computer-implemented method for controlling an aircraft, the aircraft comprising a plurality of actuators, the plurality of actuators comprising a plurality of electric propulsion units and a plurality of battery packs supplying power to the plurality of electric propulsion units, the method comprising: Determine one or more desired commands for the aircraft; At least one reference command is determined based on the desired command and one or more aircraft conditions; Monitor the energy state of the plurality of battery packs, wherein at least a first battery pack is electrically isolated from at least a second battery pack. The at least one reference command is adjusted based on the monitored state of energy of the plurality of battery packs; Generate control commands for the plurality of actuators based on at least one adjusted reference command; and The multiple actuators are controlled according to the generated control commands to satisfy one or more desired commands of the aircraft.

[0194] 12. A flight control computer, comprising: One or more memory devices, the one or more memory devices storing processor-executable instructions; and One or more processors configured to execute the instructions to cause the flight control computer to perform a computer-implemented method according to any one of clauses 1 to 11.

[0195] 13. A non-transitory computer-readable medium storing computer-readable instructions that, when executed by at least one processor, cause the at least one processor to perform a computer-implemented method according to any one of clauses 1 to 11.

[0196] 14. An aircraft comprising at least one flight control computer configured to perform a computer-implemented method according to any one of clauses 1 to 11.

[0197] 15. A flight control system, comprising: At least one memory, wherein the at least one memory stores instructions; and At least one processor, the at least one processor being configured to execute the instructions to perform a computer-implemented method according to any one of clauses 1 to 11.

[0198] 16. A flight control system for an aircraft, comprising: At least one memory, wherein the at least one memory stores instructions; and At least one processor, configured to execute the instructions to perform one or more operations, the operations including: Determine one or more desired commands for the aircraft; At least one reference command is determined based on the one or more desired commands and one or more aircraft conditions; Monitor the energy status of multiple battery packs of the aircraft, wherein at least a first battery pack of the multiple battery packs is electrically isolated from at least a second battery pack of the multiple battery packs; The at least one reference command is adjusted based on the monitored state of energy of the plurality of battery packs; Based on at least one adjusted reference command, control commands for multiple actuators of the aircraft are generated; and The multiple actuators are controlled according to the generated control commands to satisfy one or more desired commands of the aircraft.

[0199] 17. The flight control system according to Clause 16, wherein the monitored energy state includes the usable energy of the plurality of battery packs.

[0200] 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.

[0201] 19. The flight control system according to any one of Clauses 16 to 18, wherein the monitored energy state includes the difference in energy state between at least two of the plurality of battery packs.

[0202] 20. A flight control system according to any one of Clauses 16 to 19, wherein the monitored energy state includes the difference in power consumption between at least the first engine and the second engine of the plurality of electric propulsion units of the aircraft.

[0203] 21. The flight control system according to Clause 20, wherein the first engine has lower power consumption than the second engine, and wherein the generated control command causes a reduction in the power consumption of the second engine.

[0204] 22. A flight control system according to any one of clauses 16 to 21, wherein the adjustment of the at least one reference command is further based on the state of one or more engines of the aircraft's plurality of electric propulsion units.

[0205] 23. The flight control system according to any one of Clauses 16 to 22, wherein adjusting the at least one reference command comprises updating one or more reference commands to optimize the remaining discharge time across one or more high-voltage channels.

[0206] 24. The flight control system according to any one of Clauses 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.

[0207] 25. A flight control system according to any one of Clauses 16 to 24, wherein the one or more aircraft conditions include one or more of aircraft dynamics, flight conditions, or the condition of at least one aircraft component.

[0208] 26. A non-transitory computer-readable medium storing one or more instructions, said one or more instructions causing said at least one processor to perform an operation when executed by said at least one processor, said operation comprising: Determine one or more desired commands for the aircraft; At least one reference command is determined based on the one or more desired commands and one or more aircraft conditions; Monitor the energy status of multiple battery packs of the aircraft, wherein at least a first battery pack of the multiple battery packs is electrically isolated from at least a second battery pack of the multiple battery packs; The at least one reference command is adjusted based on the monitored state of energy of the plurality of battery packs; Based on at least one adjusted reference command, control commands for multiple actuators of the aircraft are generated; and The multiple actuators are controlled according to the generated control commands to satisfy one or more desired commands of the aircraft.

[0209] 27. A computer-readable medium storing computer-readable instructions that, when executed by at least one processor, cause the at least one processor to perform the method according to any one of clauses 1 to 11.

[0210] 28. A computer-implemented method, comprising: Determine one or more desired commands for the aircraft; Receive engine information for at least one of a plurality of electric propulsion units (EPUs) for 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 actuators of the aircraft; and The multiple actuators are controlled according to the generated control commands to satisfy one or more desired commands of the aircraft.

[0211] 29. The computer-implemented method according to Clause 28, wherein the engine information further includes time related to at least one temperature generated by the at least one EPU.

[0212] 30. The computer-implemented method according to Clause 29, wherein the time is generated by determining the current temperature of each of a plurality of engine components for at least one EPU.

[0213] 31. The computer-implemented method according to Clause 30, wherein the time is generated by normalizing the determined temperatures of the plurality of engine components using a time-based temperature metric, wherein one or more engine components have different temperature limits.

[0214] 32. The computer-implemented method according to 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.

[0215] 33. A computer-implemented method according to any one of clauses 28 to 32, wherein the engine information further includes the status of at least one engine associated with the plurality of EPUs.

[0216] 34. The computer-implemented method according to Clause 33, wherein the state includes one of being active or inactive.

[0217] 35. A computer-implemented method according to any one of clauses 28 to 34, wherein the generation of the control command is further based on the maximum torque that the engine can produce under normal operating conditions.

[0218] 36. A computer-implemented method according to any one of clauses 28 to 35, wherein the generation of the control command is further based on one or more aircraft conditions.

[0219] 37. The computer-implemented method according to Clause 36, wherein the one or more aircraft conditions include one or more of aircraft dynamics, flight conditions, or the condition of at least one aircraft component.

[0220] 38. The computer-implemented method according to Clause 29, wherein the time includes the remaining time determined by the machine learning model of the engine.

[0221] 39. A computer-implemented method according to any one of clauses 28 to 38, wherein the engine information further includes a temperature associated with the engine of the at least one EPU.

[0222] 40. A method for controlling an aircraft, the aircraft comprising a plurality of actuators, the plurality of actuators comprising a plurality of electric propulsion units (EPUs) and a plurality of battery packs supplying power to the plurality of electric propulsion units, the method comprising: Determine one or more desired commands for the aircraft; Receive engine information for 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 are generated for the plurality of actuators; and The multiple actuators are controlled according to the generated control commands to satisfy one or more desired commands of the aircraft.

[0223] 41. A flight control computer, comprising: One or more memory devices, the one or more memory devices storing processor-executable instructions; and One or more processors configured to execute the instructions to cause the flight control computer to perform a computer-implemented method according to any one of clauses 28 to 40.

[0224] 42. A non-transitory computer-readable medium storing computer-readable instructions that, when executed by at least one processor, cause the at least one processor to perform a computer-implemented method according to any one of clauses 28 to 40.

[0225] 43. An aircraft comprising at least one flight control computer configured to perform a computer-implemented method according to any one of clauses 28 to 40.

[0226] 44. A flight control system, comprising: At least one memory, wherein the at least one memory stores instructions; and At least one processor, the at least one processor being configured to execute the instructions to perform a computer-implemented method according to any one of clauses 28 to 40.

[0227] 45. A flight control system for an aircraft, comprising: At least one memory, wherein the at least one memory stores instructions; and At least one processor, configured to execute the instructions to perform one or more operations, the operations including: Determine one or more desired commands for the aircraft; Receive engine information for at least one of a plurality of electric propulsion units (EPUs) for 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 actuators of the aircraft; and The multiple actuators are controlled according to the generated control commands to satisfy one or more desired commands of the aircraft.

[0228] 46. ​​The flight control system according to Clause 45, wherein the engine information further includes time related to at least one temperature generated by the at least one EPU.

[0229] 47. The flight control system according to Clause 46, wherein the time is generated by determining the current temperature of each of a plurality of engine components for at least one EPU.

[0230] 48. The flight control system according to Clause 47, wherein the time is generated by normalizing the determined temperatures of the plurality of engine components using a time-based temperature metric, wherein one or more engine components have different temperature limits.

[0231] 49. The flight control system according to Clause 48, 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.

[0232] 50. The flight control system according to any one of clauses 45 to 49, wherein the engine information further includes the status of at least one engine associated with the plurality of EPUs.

[0233] 51. The flight control system according to Clause 50, wherein the state includes either active or inactive.

[0234] 52. The flight control system according to any one of clauses 45 to 51, wherein the generation of the control command is further based on the maximum torque that the engine can produce under normal operating conditions.

[0235] 53. The flight control system according to any one of clauses 45 to 52, wherein the generation of the control command is further based on one or more aircraft conditions.

[0236] 54. The flight control system pursuant to Clause 53, wherein the one or more aircraft conditions comprise one or more of aircraft dynamics, flight conditions, or the condition of at least one aircraft component.

[0237] 55. The flight control system as described in Clause 46, wherein the time includes the remaining time determined by a machine learning model of the engine.

[0238] 56. The flight control system according to any one of clauses 45 to 55, wherein the engine information further includes the temperature associated with the engine of the at least one EPU.

[0239] 57. A non-transitory computer-readable medium storing one or more instructions, said one or more instructions causing said at least one processor to perform an operation when executed by said at least one processor, said operation comprising: Determine one or more desired commands for the aircraft; Receive engine information for at least one of a plurality of electric propulsion units (EPUs) for 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 actuators of the aircraft; and The multiple actuators are controlled according to the generated control commands to satisfy one or more desired commands of the aircraft.

[0240] 58. A computer-readable medium storing computer-readable instructions that, when executed by at least one processor, cause the at least one processor to perform the method according to any one of clauses 28 to 40.

[0241] 59. An aircraft engine, comprising: At least one memory, wherein the at least one memory stores instructions; and At least one processor, configured to execute the instructions to perform one or more operations, the operations including: Determine one or more temperatures associated with one or more components of the engine; and The remaining time of the engine at the current power setting is estimated based on one or more determined temperatures, wherein the estimated remaining time corresponds to a prediction of when the engine will reach one or more predetermined limits.

[0242] 60. A computer-implemented method, comprising: Determine one or more desired commands for electric aircraft; Receive engine information for at least one engine of at least one electric propulsion unit (EPU); Receive battery information for at least one battery pack; Based on the received engine and battery information, control commands are generated for multiple actuators of the electric aircraft; and The multiple actuators are controlled according to the generated control commands to satisfy one or more desired commands of the electric aircraft.

[0243] The foregoing description has been presented for illustrative purposes. This description is not exhaustive and does not limit the invention to the precise forms or embodiments disclosed herein. Modifications and adaptations to the invention will be apparent to those skilled in the art upon consideration of the specification and practice of the disclosed embodiments of the invention.

[0244] 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 according to 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. It should also be noted that in some alternative implementations, the functions mentioned in the blocks may not appear in the order shown in the figures. For example, two blocks shown consecutively may actually be executed substantially in parallel, or the blocks may sometimes be executed in reverse order, depending on the functions involved. It should also be noted that each block in the block diagram or flowchart illustrations, and combinations of blocks in the block diagram or flowchart illustrations, may be implemented by a dedicated hardware-based system that performs the specified function or action, or a combination of dedicated hardware and computer instructions.

[0245] The foregoing description of exemplary embodiments, using flowcharts or block diagrams of methods, apparatus (systems), and computer program products, is provided with reference to such descriptions. It should be understood that each block of the flowchart or block diagram, and combinations of blocks in the flowchart or block diagram, can be implemented by the computer program product or instructions on the computer program product. These computer program instructions can be provided to a processor of a computer or other programmable data processing apparatus to produce a machine, such that the instructions, which execute via the processor of the computer or other programmable data processing apparatus, create means for implementing the functions / actions specified in one or more blocks of the flowchart or block diagram.

[0246] These computer program instructions may also be stored in a computer-readable medium that may instruct one or more hardware processors of a computer, other programmable data processing devices or other means to function in a particular manner, such that the instructions stored in the computer-readable medium form an article of writing comprising instructions that implement functions / actions specified in one or more blocks of a flowchart or block diagram.

[0247] Computer program instructions may also be loaded onto a computer, other programmable data processing apparatus or other means to cause a series of operational steps to be performed (e.g., carried out) on the computer, other programmable apparatus or other means to produce a computer-implemented process, such that the instructions, which execute on the computer or other programmable apparatus, provide for implementing a process for carrying out a function / action specified in one or more boxes of a flowchart or block diagram.

[0248] Any combination of one or more computer-readable storage media may be used. A computer-readable medium may be a non-transitory computer-readable storage medium. In the context of this document, a computer-readable storage medium can be any tangible medium that can contain or store a program for use by or in conjunction with an instruction execution system, device, or apparatus.

[0249] Program code embodied on a computer-readable medium may be transmitted using any suitable medium, including but not limited to wireless, wired, fiber optic cable, RF, IR, or any suitable combination thereof.

[0250] The features and advantages of this disclosure are apparent from the detailed description, and therefore the appended claims cover all systems and methods that fall within the true spirit and scope of this disclosure. As used herein, the indefinite articles “a” and “an” mean “one or more”. Similarly, unless the use of plural terms is explicit in the given context, it does not necessarily mean plural. Unless otherwise expressly stated, words such as “and” or “or” mean “and / or”. As used herein, unless otherwise expressly stated, “based on” may include dependent on, interdependent on, associated with, at least partially defined by, affected by, or in response to. As used herein, “related to” may include including, represented by, indicated by, or based on. Furthermore, since many modifications and alterations will be readily made by studying this disclosure, it is not intended to limit this disclosure to the exact constructions and operations shown and described, and therefore all suitable modifications and equivalents are applicable to and fall within the scope of this disclosure.

[0251] Other embodiments will be apparent to those skilled in the art upon consideration of the description and practice of the embodiments disclosed herein. The structures and circuit arrangements shown in the figures are intended for illustrative purposes only and are not intended to limit the specific arrangements and circuit configurations described and shown in the figures. The description and examples are intended to be considered merely exemplary, and the true scope and spirit of the invention are indicated by the appended claims. The foregoing description has been presented for illustrative purposes. This description is not exhaustive and does not limit the invention to the precise forms or embodiments disclosed. Modifications and adaptations to the invention will be apparent to those skilled in the art upon consideration of the description and practice of the disclosed embodiments of the invention.

Claims

1. A computer-implemented method, comprising: Determine one or more desired commands for the aircraft; At least one reference command is determined based on the one or more desired commands and one or more aircraft conditions; Monitor the energy status of multiple battery packs of the aircraft, wherein at least a first battery pack of the multiple battery packs is electrically isolated from at least a second battery pack of the multiple battery packs; The at least one reference command is adjusted based on the monitored state of energy of the plurality of battery packs; Based on at least one adjusted reference command, control commands for multiple actuators of the aircraft are generated. as well as The multiple actuators are controlled according to the generated control commands to satisfy one or more desired commands of the aircraft.

2. The computer-implemented method of claim 1, wherein the monitored state of energy includes the usable energy of the plurality of battery packs.

3. The computer-implemented method 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. The computer-implemented method according to any one of claims 1 to 3, wherein the monitored state of energy includes the difference in state of energy between at least two of the plurality of battery packs.

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

6. The computer-implemented method of claim 5, wherein the first engine has lower power consumption than the second engine, and wherein the generated control command causes a reduction in the power consumption of the second engine.

7. The computer-implemented method according to any one of claims 1 to 6, wherein the adjustment of the at least one reference command is further based on the state of one or more engines of the plurality of electric propulsion units of the aircraft.

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

9. The computer-implemented method 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 computer-implemented method according to any one of claims 1 to 9, wherein the one or more aircraft states comprise one or more of aircraft dynamics, flight status, or the state of at least one aircraft component.

11. A computer-implemented method for controlling an aircraft, the aircraft comprising a plurality of actuators, the plurality of actuators comprising a plurality of electric propulsion units and a plurality of battery packs supplying power to the plurality of electric propulsion units, the method comprising: Determine one or more desired commands for the aircraft; At least one reference command is determined based on the desired command and one or more aircraft conditions; Monitor the energy state of the plurality of battery packs, wherein at least a first battery pack is electrically isolated from at least a second battery pack. The at least one reference command is adjusted based on the monitored state of energy of the plurality of battery packs; Control commands for the plurality of actuators are generated based on at least one adjusted reference command. as well as The multiple actuators are controlled according to the generated control commands to satisfy one or more desired commands of the aircraft.

12. A flight control computer, comprising: One or more memory devices storing processor-executable instructions; as well as One or more processors configured to execute the instructions to cause the flight control computer to perform a computer-implemented method according to any one of claims 1 to 11.

13. A non-transitory computer-readable medium storing computer-readable instructions that, when executed by at least one processor, cause the at least one processor to perform a computer-implemented method according to any one of claims 1 to 11.

14. An aircraft comprising at least one flight control computer configured to perform a computer-implemented method according to any one of claims 1 to 11.

15. A flight control system, comprising: At least one memory, wherein the at least one memory stores instructions; as well as At least one processor, the at least one processor being configured to execute the instructions to perform a computer-implemented method according to any one of claims 1 to 11.

16. A flight control system for an aircraft, comprising: At least one memory, wherein the at least one memory stores instructions; as well as At least one processor, configured to execute the instructions to perform one or more operations, the operations including: Determine one or more desired commands for the aircraft; At least one reference command is determined based on the one or more desired commands and one or more aircraft conditions; Monitor the energy status of multiple battery packs of the aircraft, wherein at least a first battery pack of the multiple battery packs is electrically isolated from at least a second battery pack of the multiple battery packs; The at least one reference command is adjusted based on the monitored state of energy of the plurality of battery packs; Based on at least one adjusted reference command, control commands for multiple actuators of the aircraft are generated; and The multiple actuators are controlled according to the generated control commands to satisfy one or more desired commands of the aircraft.

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

18. The flight control system of 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 state 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 the first engine and the second engine of the plurality of electric propulsion units of the aircraft.

21. The flight control system of claim 20, wherein the first engine has lower power consumption than the second engine, and wherein the generated control command causes a reduction in the power consumption of the second engine.

22. The flight control system according to any one of claims 16 to 21, wherein the adjustment of the at least one reference command is further based on the state of one or more engines of the aircraft's plurality of 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 comprises 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 states include one or more of aircraft dynamics, flight status, or the state of at least one aircraft component.

26. A non-transitory computer-readable medium storing one or more instructions, said one or more instructions causing said at least one processor to perform an operation when executed by said at least one processor, said operation comprising: Determine one or more desired commands for the aircraft; At least one reference command is determined based on the one or more desired commands and one or more aircraft conditions; Monitor the energy status of multiple battery packs of the aircraft, wherein at least a first battery pack of the multiple battery packs is electrically isolated from at least a second battery pack of the multiple battery packs; The at least one reference command is adjusted based on the monitored state of energy of the plurality of battery packs; Based on at least one adjusted reference command, control commands for multiple actuators of the aircraft are generated. as well as The multiple actuators are controlled according to the generated control commands to satisfy one or more desired commands of the aircraft.

27. A computer-readable medium storing computer-readable instructions that, when executed by at least one processor, cause the at least one processor to perform the method according to any one of claims 1 to 11.

28. A computer-implemented method, comprising: Determine one or more desired commands for the aircraft; Receive engine information for at least one of a plurality of electric propulsion units (EPUs) for 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 actuators of the aircraft; as well as The multiple actuators are controlled according to the generated control commands to satisfy one or more desired commands of the aircraft.

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

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

31. The computer-implemented method of claim 30, wherein the time is generated by normalizing the determined temperatures of the plurality of engine components using a time-based temperature metric, wherein one or more engine components have different temperature limits.

32. The computer-implemented method of 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 computer-implemented method 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 computer-implemented method of claim 33, wherein the state includes one of being active or inactive.

35. The computer-implemented method 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 produce under normal operating conditions.

36. The computer-implemented method according to any one of claims 28 to 35, wherein the generation of the control command is further based on one or more aircraft conditions.

37. The computer-implemented method of claim 36, wherein the one or more aircraft states comprise one or more of aircraft dynamics, flight status, or the state of at least one aircraft component.

38. The computer-implemented method of claim 29, wherein the time includes the remaining time determined by the machine learning model of the engine.

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

40. A method for controlling an aircraft, the aircraft comprising a plurality of actuators, the plurality of actuators comprising a plurality of electric propulsion units (EPUs) and a plurality of battery packs supplying power to the plurality of electric propulsion units, the method comprising: Determine one or more desired commands for the aircraft; Receive engine information for 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 actuators are generated; as well as The multiple actuators are controlled according to the generated control commands to satisfy one or more desired commands of the aircraft.

41. A flight control computer, comprising: One or more memory devices storing processor-executable instructions; as well as One or more processors configured to execute the instructions to cause the flight control computer to perform the computer-implemented method according to any one of claims 28 to 40.

42. A non-transitory computer-readable medium storing computer-readable instructions that, when executed by at least one processor, cause the at least one processor to perform a computer-implemented method according to any one of claims 28 to 40.

43. An aircraft comprising at least one flight control computer configured to perform a computer-implemented method according to any one of claims 28 to 40.

44. A flight control system, comprising: At least one memory, wherein the at least one memory stores instructions; as well as At least one processor, the at least one processor being configured to execute the instructions to perform a computer-implemented method according to any one of claims 28 to 40.

45. A flight control system for an aircraft, comprising: At least one memory, wherein the at least one memory stores instructions; as well as At least one processor, configured to execute the instructions to perform one or more operations, the operations including: Determine one or more desired commands for the aircraft; Receive engine information for at least one of a plurality of electric propulsion units (EPUs) for 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 actuators of the aircraft; and The multiple actuators are controlled according to the generated control commands to satisfy one or more desired commands of the aircraft.

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

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

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

49. The flight control system of claim 48, 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.

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 of claim 50, wherein the state includes either active or inactive.

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

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

54. The flight control system of claim 53, wherein the one or more aircraft states comprise one or more of aircraft dynamics, flight status, or the state of at least one aircraft component.

55. The flight control system of claim 46, wherein the time includes the remaining time determined by a 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-transitory computer-readable medium storing one or more instructions, said one or more instructions causing said at least one processor to perform an operation when executed by said at least one processor, said operation comprising: Determine one or more desired commands for the aircraft; Receive engine information for at least one of a plurality of electric propulsion units (EPUs) for 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 actuators of the aircraft; as well as The multiple actuators are controlled according to the generated control commands to satisfy one or more desired commands of the aircraft.

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

59. An aircraft engine, comprising: At least one memory, wherein the at least one memory stores instructions; as well as At least one processor, configured to execute the instructions to perform one or more operations, the operations including: Determine one or more temperatures associated with one or more components of the engine; as well as The remaining time of the engine at the current power setting is estimated based on one or more determined temperatures, wherein the estimated remaining time corresponds to a prediction of when the engine will reach one or more predetermined limits.

60. A computer-implemented method, comprising: Determine one or more desired commands for electric aircraft; Receive engine information for at least one engine of at least one electric propulsion unit (EPU); Receive battery information for at least one battery pack; Based on the received engine and battery information, control commands are generated for multiple actuators of the electric aircraft. as well as The multiple actuators are controlled according to the generated control commands to satisfy one or more desired commands of the electric aircraft.

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