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WO2025015072A1 - Systèmes et procédés pour protections dynamiques de canaux de puissance - Google Patents

Systèmes et procédés pour protections dynamiques de canaux de puissance Download PDF

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Publication number
WO2025015072A1
WO2025015072A1 PCT/US2024/037420 US2024037420W WO2025015072A1 WO 2025015072 A1 WO2025015072 A1 WO 2025015072A1 US 2024037420 W US2024037420 W US 2024037420W WO 2025015072 A1 WO2025015072 A1 WO 2025015072A1
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WO
WIPO (PCT)
Prior art keywords
power
aircraft
voltage bus
limit
engine
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Pending
Application number
PCT/US2024/037420
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English (en)
Inventor
Nathan Thomas Depenbusch
Alessandra CARNO
Alfonso Noriega BENEKE
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Archer Aviation Inc
Original Assignee
Archer Aviation Inc
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Archer Aviation Inc filed Critical Archer Aviation Inc
Priority to PCT/US2024/037420 priority Critical patent/WO2025015072A1/fr
Publication of WO2025015072A1 publication Critical patent/WO2025015072A1/fr
Pending legal-status Critical Current
Anticipated expiration legal-status Critical

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Classifications

    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
    • G05D1/00—Control of position, course, altitude or attitude of land, water, air or space vehicles, e.g. using automatic pilots
    • G05D1/20—Control system inputs
    • G05D1/24—Arrangements for determining position or orientation
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B60—VEHICLES IN GENERAL
    • B60L—PROPULSION OF ELECTRICALLY-PROPELLED VEHICLES; SUPPLYING ELECTRIC POWER FOR AUXILIARY EQUIPMENT OF ELECTRICALLY-PROPELLED VEHICLES; ELECTRODYNAMIC BRAKE SYSTEMS FOR VEHICLES IN GENERAL; MAGNETIC SUSPENSION OR LEVITATION FOR VEHICLES; MONITORING OPERATING VARIABLES OF ELECTRICALLY-PROPELLED VEHICLES; ELECTRIC SAFETY DEVICES FOR ELECTRICALLY-PROPELLED VEHICLES
    • B60L3/00—Electric devices on electrically-propelled vehicles for safety purposes; Monitoring operating variables, e.g. speed, deceleration or energy consumption
    • B60L3/0023—Detecting, eliminating, remedying or compensating for drive train abnormalities, e.g. failures within the drive train
    • B60L3/0038—Detecting, eliminating, remedying or compensating for drive train abnormalities, e.g. failures within the drive train relating to sensors
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B60—VEHICLES IN GENERAL
    • B60L—PROPULSION OF ELECTRICALLY-PROPELLED VEHICLES; SUPPLYING ELECTRIC POWER FOR AUXILIARY EQUIPMENT OF ELECTRICALLY-PROPELLED VEHICLES; ELECTRODYNAMIC BRAKE SYSTEMS FOR VEHICLES IN GENERAL; MAGNETIC SUSPENSION OR LEVITATION FOR VEHICLES; MONITORING OPERATING VARIABLES OF ELECTRICALLY-PROPELLED VEHICLES; ELECTRIC SAFETY DEVICES FOR ELECTRICALLY-PROPELLED VEHICLES
    • B60L3/00—Electric devices on electrically-propelled vehicles for safety purposes; Monitoring operating variables, e.g. speed, deceleration or energy consumption
    • B60L3/0023—Detecting, eliminating, remedying or compensating for drive train abnormalities, e.g. failures within the drive train
    • B60L3/0084—Detecting, eliminating, remedying or compensating for drive train abnormalities, e.g. failures within the drive train relating to control modules
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B60—VEHICLES IN GENERAL
    • B60L—PROPULSION OF ELECTRICALLY-PROPELLED VEHICLES; SUPPLYING ELECTRIC POWER FOR AUXILIARY EQUIPMENT OF ELECTRICALLY-PROPELLED VEHICLES; ELECTRODYNAMIC BRAKE SYSTEMS FOR VEHICLES IN GENERAL; MAGNETIC SUSPENSION OR LEVITATION FOR VEHICLES; MONITORING OPERATING VARIABLES OF ELECTRICALLY-PROPELLED VEHICLES; ELECTRIC SAFETY DEVICES FOR ELECTRICALLY-PROPELLED VEHICLES
    • B60L58/00—Methods or circuit arrangements for monitoring or controlling batteries or fuel cells, specially adapted for electric vehicles
    • B60L58/10—Methods or circuit arrangements for monitoring or controlling batteries or fuel cells, specially adapted for electric vehicles for monitoring or controlling batteries
    • B60L58/12—Methods or circuit arrangements for monitoring or controlling batteries or fuel cells, specially adapted for electric vehicles for monitoring or controlling batteries responding to state of charge [SoC]
    • B60L58/14—Preventing excessive discharging
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B60—VEHICLES IN GENERAL
    • B60L—PROPULSION OF ELECTRICALLY-PROPELLED VEHICLES; SUPPLYING ELECTRIC POWER FOR AUXILIARY EQUIPMENT OF ELECTRICALLY-PROPELLED VEHICLES; ELECTRODYNAMIC BRAKE SYSTEMS FOR VEHICLES IN GENERAL; MAGNETIC SUSPENSION OR LEVITATION FOR VEHICLES; MONITORING OPERATING VARIABLES OF ELECTRICALLY-PROPELLED VEHICLES; ELECTRIC SAFETY DEVICES FOR ELECTRICALLY-PROPELLED VEHICLES
    • B60L58/00—Methods or circuit arrangements for monitoring or controlling batteries or fuel cells, specially adapted for electric vehicles
    • B60L58/10—Methods or circuit arrangements for monitoring or controlling batteries or fuel cells, specially adapted for electric vehicles for monitoring or controlling batteries
    • B60L58/12—Methods or circuit arrangements for monitoring or controlling batteries or fuel cells, specially adapted for electric vehicles for monitoring or controlling batteries responding to state of charge [SoC]
    • B60L58/15—Preventing overcharging
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B60—VEHICLES IN GENERAL
    • B60L—PROPULSION OF ELECTRICALLY-PROPELLED VEHICLES; SUPPLYING ELECTRIC POWER FOR AUXILIARY EQUIPMENT OF ELECTRICALLY-PROPELLED VEHICLES; ELECTRODYNAMIC BRAKE SYSTEMS FOR VEHICLES IN GENERAL; MAGNETIC SUSPENSION OR LEVITATION FOR VEHICLES; MONITORING OPERATING VARIABLES OF ELECTRICALLY-PROPELLED VEHICLES; ELECTRIC SAFETY DEVICES FOR ELECTRICALLY-PROPELLED VEHICLES
    • B60L58/00—Methods or circuit arrangements for monitoring or controlling batteries or fuel cells, specially adapted for electric vehicles
    • B60L58/10—Methods or circuit arrangements for monitoring or controlling batteries or fuel cells, specially adapted for electric vehicles for monitoring or controlling batteries
    • B60L58/16—Methods or circuit arrangements for monitoring or controlling batteries or fuel cells, specially adapted for electric vehicles for monitoring or controlling batteries responding to battery ageing, e.g. to the number of charging cycles or the state of health [SoH]
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B60—VEHICLES IN GENERAL
    • B60L—PROPULSION OF ELECTRICALLY-PROPELLED VEHICLES; SUPPLYING ELECTRIC POWER FOR AUXILIARY EQUIPMENT OF ELECTRICALLY-PROPELLED VEHICLES; ELECTRODYNAMIC BRAKE SYSTEMS FOR VEHICLES IN GENERAL; MAGNETIC SUSPENSION OR LEVITATION FOR VEHICLES; MONITORING OPERATING VARIABLES OF ELECTRICALLY-PROPELLED VEHICLES; ELECTRIC SAFETY DEVICES FOR ELECTRICALLY-PROPELLED VEHICLES
    • B60L58/00—Methods or circuit arrangements for monitoring or controlling batteries or fuel cells, specially adapted for electric vehicles
    • B60L58/10—Methods or circuit arrangements for monitoring or controlling batteries or fuel cells, specially adapted for electric vehicles for monitoring or controlling batteries
    • B60L58/24—Methods or circuit arrangements for monitoring or controlling batteries or fuel cells, specially adapted for electric vehicles for monitoring or controlling batteries for controlling the temperature of batteries
    • B60L58/25—Methods or circuit arrangements for monitoring or controlling batteries or fuel cells, specially adapted for electric vehicles for monitoring or controlling batteries for controlling the temperature of batteries by controlling the electric load
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C29/00—Aircraft capable of landing or taking-off vertically, e.g. vertical take-off and landing [VTOL] aircraft
    • B64C29/0008—Aircraft capable of landing or taking-off vertically, e.g. vertical take-off and landing [VTOL] aircraft having its flight directional axis horizontal when grounded
    • B64C29/0016—Aircraft capable of landing or taking-off vertically, e.g. vertical take-off and landing [VTOL] aircraft having its flight directional axis horizontal when grounded the lift during taking-off being created by free or ducted propellers or by blowers
    • B64C29/0025—Aircraft capable of landing or taking-off vertically, e.g. vertical take-off and landing [VTOL] aircraft having its flight directional axis horizontal when grounded the lift during taking-off being created by free or ducted propellers or by blowers the propellers being fixed relative to the fuselage
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64U—UNMANNED AERIAL VEHICLES [UAV]; EQUIPMENT THEREFOR
    • B64U50/00—Propulsion; Power supply
    • B64U50/10—Propulsion
    • B64U50/19—Propulsion using electrically powered motors
    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
    • G05D1/00—Control of position, course, altitude or attitude of land, water, air or space vehicles, e.g. using automatic pilots
    • G05D1/40—Control within particular dimensions
    • G05D1/48—Control of altitude or depth
    • G05D1/485—Control of rate of change of altitude or depth
    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
    • G05D1/00—Control of position, course, altitude or attitude of land, water, air or space vehicles, e.g. using automatic pilots
    • G05D1/80—Arrangements for reacting to or preventing system or operator failure
    • G05D1/82—Limited authority control, e.g. enforcing a flight envelope
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B60—VEHICLES IN GENERAL
    • B60L—PROPULSION OF ELECTRICALLY-PROPELLED VEHICLES; SUPPLYING ELECTRIC POWER FOR AUXILIARY EQUIPMENT OF ELECTRICALLY-PROPELLED VEHICLES; ELECTRODYNAMIC BRAKE SYSTEMS FOR VEHICLES IN GENERAL; MAGNETIC SUSPENSION OR LEVITATION FOR VEHICLES; MONITORING OPERATING VARIABLES OF ELECTRICALLY-PROPELLED VEHICLES; ELECTRIC SAFETY DEVICES FOR ELECTRICALLY-PROPELLED VEHICLES
    • B60L2200/00—Type of vehicles
    • B60L2200/10—Air crafts
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B60—VEHICLES IN GENERAL
    • B60L—PROPULSION OF ELECTRICALLY-PROPELLED VEHICLES; SUPPLYING ELECTRIC POWER FOR AUXILIARY EQUIPMENT OF ELECTRICALLY-PROPELLED VEHICLES; ELECTRODYNAMIC BRAKE SYSTEMS FOR VEHICLES IN GENERAL; MAGNETIC SUSPENSION OR LEVITATION FOR VEHICLES; MONITORING OPERATING VARIABLES OF ELECTRICALLY-PROPELLED VEHICLES; ELECTRIC SAFETY DEVICES FOR ELECTRICALLY-PROPELLED VEHICLES
    • B60L2220/00—Electrical machine types; Structures or applications thereof
    • B60L2220/40—Electrical machine applications
    • B60L2220/42—Electrical machine applications with use of more than one motor
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B60—VEHICLES IN GENERAL
    • B60L—PROPULSION OF ELECTRICALLY-PROPELLED VEHICLES; SUPPLYING ELECTRIC POWER FOR AUXILIARY EQUIPMENT OF ELECTRICALLY-PROPELLED VEHICLES; ELECTRODYNAMIC BRAKE SYSTEMS FOR VEHICLES IN GENERAL; MAGNETIC SUSPENSION OR LEVITATION FOR VEHICLES; MONITORING OPERATING VARIABLES OF ELECTRICALLY-PROPELLED VEHICLES; ELECTRIC SAFETY DEVICES FOR ELECTRICALLY-PROPELLED VEHICLES
    • B60L2240/00—Control parameters of input or output; Target parameters
    • B60L2240/40—Drive Train control parameters
    • B60L2240/54—Drive Train control parameters related to batteries
    • B60L2240/545—Temperature
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B60—VEHICLES IN GENERAL
    • B60L—PROPULSION OF ELECTRICALLY-PROPELLED VEHICLES; SUPPLYING ELECTRIC POWER FOR AUXILIARY EQUIPMENT OF ELECTRICALLY-PROPELLED VEHICLES; ELECTRODYNAMIC BRAKE SYSTEMS FOR VEHICLES IN GENERAL; MAGNETIC SUSPENSION OR LEVITATION FOR VEHICLES; MONITORING OPERATING VARIABLES OF ELECTRICALLY-PROPELLED VEHICLES; ELECTRIC SAFETY DEVICES FOR ELECTRICALLY-PROPELLED VEHICLES
    • B60L2240/00—Control parameters of input or output; Target parameters
    • B60L2240/40—Drive Train control parameters
    • B60L2240/54—Drive Train control parameters related to batteries
    • B60L2240/547—Voltage
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B60—VEHICLES IN GENERAL
    • B60L—PROPULSION OF ELECTRICALLY-PROPELLED VEHICLES; SUPPLYING ELECTRIC POWER FOR AUXILIARY EQUIPMENT OF ELECTRICALLY-PROPELLED VEHICLES; ELECTRODYNAMIC BRAKE SYSTEMS FOR VEHICLES IN GENERAL; MAGNETIC SUSPENSION OR LEVITATION FOR VEHICLES; MONITORING OPERATING VARIABLES OF ELECTRICALLY-PROPELLED VEHICLES; ELECTRIC SAFETY DEVICES FOR ELECTRICALLY-PROPELLED VEHICLES
    • B60L2240/00—Control parameters of input or output; Target parameters
    • B60L2240/40—Drive Train control parameters
    • B60L2240/54—Drive Train control parameters related to batteries
    • B60L2240/549—Current
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B60—VEHICLES IN GENERAL
    • B60L—PROPULSION OF ELECTRICALLY-PROPELLED VEHICLES; SUPPLYING ELECTRIC POWER FOR AUXILIARY EQUIPMENT OF ELECTRICALLY-PROPELLED VEHICLES; ELECTRODYNAMIC BRAKE SYSTEMS FOR VEHICLES IN GENERAL; MAGNETIC SUSPENSION OR LEVITATION FOR VEHICLES; MONITORING OPERATING VARIABLES OF ELECTRICALLY-PROPELLED VEHICLES; ELECTRIC SAFETY DEVICES FOR ELECTRICALLY-PROPELLED VEHICLES
    • B60L2240/00—Control parameters of input or output; Target parameters
    • B60L2240/60—Navigation input
    • B60L2240/62—Vehicle position
    • B60L2240/622—Vehicle position by satellite navigation
    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
    • G05D2105/00—Specific applications of the controlled vehicles
    • G05D2105/20—Specific applications of the controlled vehicles for transportation
    • G05D2105/22—Specific applications of the controlled vehicles for transportation of humans
    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
    • G05D2109/00—Types of controlled vehicles
    • G05D2109/20—Aircraft, e.g. drones
    • G05D2109/22—Aircraft, e.g. drones with fixed wings
    • G05D2109/23—Vertical take-off and landing [VTOL] aircraft; Short take-off and landing [STOL, STOVL] aircraft
    • G05D2109/24—Convertible aircraft, e.g. tiltrotor aircraft
    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
    • G05D2111/00—Details of signals used for control of position, course, altitude or attitude of land, water, air or space vehicles
    • G05D2111/50—Internal signals, i.e. from sensors located in the vehicle, e.g. from compasses or angular sensors

Definitions

  • This disclosure relates generally to powered aerial vehicles. More particularly, and without limitation, the present disclosure relates to innovations in aircrafts driven by electric propulsion systems. Certain aspects of the present disclosure generally relate to systems and methods for flight control of aircrafts driven by electric propulsion systems and in other types of vehicles, as well as flight control of aircrafts in flight simulators and video games. Other aspects of the present disclosure generally relate to improvements in flight control systems and methods that provide particular advantages in aerial vehicles and may be used in other types of vehicles.
  • the inventors here have recognized several problems that may be associated with flight control of aircraft, including a tilt-rotor aircraft that uses electrical or hybrid-electric propulsion systems (hereinafter referred to as electric propulsion units or “EPUs”).
  • EPUs electric propulsion units
  • Aircraft relying of electrical of hybrid-electric propulsion systems require stricter power budgeting and monitoring to ensure reliable operation of key EPU components and for overall safe operation of the aircraft.
  • Managing power usage of an electric aircraft is of crucial importance for maintaining optimal performance and safety during aircraft operation.
  • Prioritization of power allocation to critical systems and components ensures reliable aircraft operation in particular during key phases such as takeoff, landing, and transitions, which may require increased amounts of power for safe execution, as well as in cases of faults, failures or emergencies, where the availability of power supply may change quickly and unexpectedly.
  • careful power control aids in mitigation of challenges including thermal management and prolonged recharging times, which are essential for maintaining operational reliability and achieving environmental goals electric technologies aim to address.
  • the present disclosure relates generally to flight control of electric aircraft and other powered aerial vehicles. More particularly, and without limitation, the present disclosure relates to innovations useable in tilt-rotor aircraft that use electrical propulsion systems. For example, certain aspects of the present disclosure relate to providing systems and method for limiting of power delivered from high-voltage bus to electrical components in response to power demand from said components that may exceed a power limit of a high-voltage (HV) bus.
  • HV high-voltage
  • Techniques of present disclosure may relate to control of power flow between the electrical components (e.g. engines) and an energy source (e.g. a battery pack) across various flight configurations.
  • some electrical components connected to an HV bus may demand power, while other components may generate power.
  • some propellers or rotors may be used for stabilizing and/or directing the aircraft demanding power from the battery pack.
  • at least some of the propellers of the aircraft may windmill, generating energy that may be stored in the batteries of the battery pack.
  • some electric components e.g. electric engines may function as both power consumers and generators, and in systems with multiple engines connected to the same power bus, the total power consumed or generated can exceed the power source's capacity during extreme maneuvers, which may result in equipment and/or power supply failures.
  • a control system may be required to actively balance energy demand and generation to within pre-defined and/or dynamic safety limits.
  • This control may become complex, necessitating centralized management equipped with dynamic power allocation capabilities to respond to changing flight conditions as well as potential faults and failure.
  • Some embodiments of the present disclosure may relate to a method for allocating power to components of a high-voltage bus of an aircraft.
  • the method for allocating power may comprise receiving an input indicative of an operation command and calculating, based on the received input, a total power demand associated with a plurality of components connected to the high-voltage bus, wherein each component may have a respective component power demand.
  • the method may further comprise comparing the total power demand to a high- voltage bus power limit assigned to the high-voltage bus to determine a difference between the high-voltage bus power limit and the total power demand, which may be followed by determining power allocation to each of the components connected to the high-voltage bus based on the difference between the high-voltage bus power limit and the total power demand.
  • the method may also allocate power to each component connected to the high-voltage bus using the determined power allocation.
  • the high-voltage bus may be electrically connected to a plurality of battery cells and the method for allocating power may determine the high-voltage bus power limit assigned to the high-voltage bus based on a hardware power limit and electrical power available from the plurality of battery cells.
  • allocating power may comprise reducing power allocated to all components connected to the high-voltage bus, wherein the power reduction for all components may be proportional or disproportionate. In some embodiments, allocating power may comprise reducing power allocated to at least some components connected to the high-voltage bus, wherein reducing power allocated to at least some components comprises: reducing power based on a ranking of the components connected to the high-voltage bus.
  • the components connected to the high-voltage bus comprise at least one engine and allocating the power may include modifying a command to request a reduced amount of torque from the at least one engine.
  • total power demand may be determined based on a torque demand of the at least one engine, while determining a torque limit for the at least one engine may be based on the determined power allocation.
  • the method for allocating power to components of a high-voltage bus of an aircraft may further comprise determining a difference in power demanded and power allocated for the at least one engine, determining a torque limit for the engine based on the determined difference and updating an engine reference model based on the determined torque limit. Determination of power allocation may be executed by a high-voltage protection function, and power allocation may be determined by the high-voltage protection function for every received input indicative of an operation command. In some embodiments, the high-voltage protection function may be executed by a flight control computer in the aircraft while the aircraft may be in flight. The method may be initiated in response to a fault or a failure.
  • calculating the total power demand may be based on an efficiency of each component, wherein the efficiency of the component may be measured during operation of the component or may be predicted based on a mathematical model of operation of the component. Determination of power allocation may be performed dynamically and the determined power allocation may change over time.
  • the method for allocating power to components of a high-voltage bus of an aircraft may further comprise determining that the total power demand may exceed an activation threshold, activating a power allocation function based on the determination that the total power demand may exceed the activation threshold, wherein the power allocation function is configured to determine the power allocation and deactivating the power allocation upon detection that the total power demand may be below a deactivation threshold.
  • the activation threshold is higher than the deactivation threshold.
  • Some embodiments of the present disclosure may relate to a vertical take-off and landing (VTOL) aircraft.
  • the aircraft may comprise a fuselage, at least one wings mounted to the fuselage, at least one stabilizer mounter to the rear of the fuselage and a plurality of propellers mounted to the at least one wing, wherein at least one of the propellers may be tiltable.
  • the aircraft may also comprise a high-voltage bus that may be electrically connected to a plurality of battery cells and a flight control system that may comprise at least one processing unit configured to carry out the method for allocating power to components of a high-voltage bus of an aircraft.
  • the system may comprise a high- voltage bus and at least one processor.
  • the at least one processor may be configured to receive an input indicative of an operation command, calculate, based on the received input, a total power demand associated with a plurality of components connected to the high-voltage bus, each component having a respective component power demand, compare the total power demand to a high-voltage bus power limit assigned to the high-voltage bus to determine a difference between the high-voltage bus power limit and the total power demand determine power allocation to each of the components connected to the high-voltage bus based on the difference between the high-voltage bus power limit and the total power demand and allocate power to each component connected to the high-voltage bus using the determined power allocation.
  • the high-voltage bus may be electrically connected to a plurality of battery cells.
  • the components connected to the high-voltage bus may comprise at least one engine.
  • the at least one processor may be further configure to allocate the power to each component connected to the high-voltage bus by modifying a command to request a reduced amount of torque from the at least one engine.
  • the at least one processor may also determine the total power demand based on a torque demand of the at least one engine and optionally to determine a torque limit for the at least one engine based on the determined power allocation.
  • the at least one processor may be further configured to determine a difference in power demand and power allocated for the at least one engine, determine a torque limit for the engine based on the determined difference and update an engine reference model based on the determined torque limit. In some embodiments, the at least one processor may determine that the total power demand exceeds an activation threshold, activate a power allocation function based on the determination that the total power demand exceeds the activation threshold, wherein the power allocation function is configured to determine the power allocation and deactivate the power allocation upon detection that the total power demand is below a deactivation threshold.
  • the at least one processor may calculate the total power demand based on an efficiency of each component, wherein the efficiency of the component may be measured during operation of the component or wherein the efficiency of the component may be predicted based on a mathematical model of operation of the component.
  • Fig. 1 shows an exemplary VTOL aircraft, consistent with disclosed embodiments.
  • Fig. 2 shows an exemplary VTOL aircraft, consistent with disclosed embodiments.
  • FIG. 3 shows an exemplary top plan view of a VTOL aircraft, consistent with disclosed embodiments.
  • FIG. 4 illustrates exemplary propeller rotation of a VTOL aircraft, consistent with disclosed embodiments.
  • Fig. 5 shows exemplary power connections in a VTOL aircraft, consistent with disclosed embodiments.
  • Fig. 6 shows an exemplary architecture of an electric propulsion unit, consistent with disclosed embodiments.
  • Fig. 7 shows an exemplary top plan view of a VTOL aircraft, consistent with disclosed embodiments.
  • Fig. 8 shows an exemplary flight control signaling architecture, consistent with disclosed embodiments.
  • FIGs. 9A-9F illustrate exemplary top plan views of VTOL aircraft, consistent with disclosed embodiments.
  • Fig. 10 illustrates a functional block diagram of an exemplary control system of an electric VTOL aircraft, consistent with disclosed embodiments.
  • Fig. 11 illustrates a functional block diagram of an exemplary implementation of a method for allocating power to components of high-voltage bus, according to some embodiments.
  • Fig. 12 shows an example of power demand from plurality of engines of an aircraft according to some embodiments.
  • Fig. 13 shows an example of allocation and limiting to plurality of engines of an aircraft according to some embodiments.
  • Fig. 14 shows an example of total power demand calculation for plurality of engines of an aircraft when some engines consume power and other engines generate power according to some embodiments.
  • Fig. 15A shows an example of power reduction in response to fault or failure according to some embodiments.
  • Fig. 15B shows an example of power reduction in response to fault or failure occurring in a component connected to a different HV bus of a plurality of HV buses.
  • Fig. 16. illustrates an exemplary operation of power limiting by HV protection function according to some embodiment.
  • the present disclosure addresses systems, components, and techniques primarily for use in an aircraft.
  • the aircraft may be an aircraft with a pilot, an aircraft without a pilot (e.g., a UAV), a drone, a helicopter, and/or an airplane.
  • An aircraft includes a physical body and one or more components (e.g., a wing, a tail, a propeller) configured to allow the aircraft to fly.
  • the aircraft may include any configuration that includes at least one propeller.
  • the aircraft is driven by one or more electric propulsion systems (hereinafter referred to as electric propulsion units or “EPUs”).
  • EPUs electric propulsion units
  • the aircraft may be fully electric, hybrid, or gas powered.
  • the aircraft is a tilt-rotor aircraft configured for frequent (e.g., over 50 flights per work day), short-duration flights (e.g., less than 100 miles per flight) over, into, and out of densely populated regions.
  • the aircraft may be configured to carry 4-6 passengers or commuters who have an expectation of a comfortable experience with low noise and low vibration. Accordingly, it is desirable to control the power allocation to components of the aircraft in a manner that increases energy efficiency to improve aircraft performance (e.g., increase safety, increase flight range, reduce aircraft weight, increase fuel efficiency, increase ride comfort, increase payload capacity, and/or increase structural integrity).
  • Disclosed embodiments provide new and improved configurations of aircraft components, some of which are not observed in conventional aircraft, and/or identified design criteria for components that differ from those of conventional aircraft. Such alternate configurations and design criteria, in combination addressing drawbacks and challenges with conventional components, yielded the embodiments disclosed herein for various configurations and designs of components for an aircraft (e.g., electric aircraft or hybrid-electric aircraft) driven by a propulsion system.
  • an aircraft e.g., electric aircraft or hybrid-electric aircraft driven by a propulsion system.
  • the aircraft driven by a propulsion system of the present disclosure may be designed to be capable of both vertical and conventional takeoff and landing, with a distributed propulsion system enabling vertical flight, horizontal and lateral flight, and transition (e.g., transitioning between vertical flight and horizontal flight).
  • the aircraft may generate thrust by supplying high voltage electrical power to a plurality of engines of the distributed propulsion system, which may include components to convert the high voltage electrical power into mechanical shaft power to rotate a propeller.
  • Embodiments may include an electric engine (e.g., motor) connected to an onboard electrical power source, which may include a device capable of storing energy such as a battery or capacitor, and may optionally include one or more systems for harnessing or generating electricity such as a fuel powered generator or solar panel array.
  • the aircraft may comprise a hybrid aircraft configured to use at least one of an electric-based energy source or a fuel-based energy source to power the distributed propulsion system.
  • the aircraft may be powered by one or more batteries, internal combustion engines (ICE), generators, turbine engines, or ducted fans.
  • ICE internal combustion engines
  • the engines may be mounted directly to the wing, or mounted to one or more booms attached to the wing.
  • the amount of thrust each engine generates may be governed by a torque command from a Flight Control System (FCS) over a digital communication interface to each engine.
  • FCS Flight Control System
  • Embodiments may include forward engines (and associated propellers) that are capable of altering their orientation, or tilt.
  • the engines may rotate the propellers in a clockwise or counterclockwise direction.
  • the difference in propeller rotation direction may be achieved using the direction of engine rotation.
  • the engines may all rotate in the same direction, and gearing may be used to achieve different propeller rotation directions.
  • an aircraft may possess quantities of engines in various combinations of forward and aft engine configurations.
  • a forward engine may be considered an engine that is positioned predominantly towards the leading edge of a wing.
  • An aft engine may be considered an engine that is positioned predominantly towards the trailing edge of a wing.
  • an aircraft may possess six forward and six aft engines, five forward and five aft engines, four forward and four aft engines, three forward and three aft engines, two forward and two aft engines, or any other combination of forward and aft engines, including embodiments where the number of forward engines and aft engines are not equivalent.
  • the forward and aft engines may provide vertical thrust during takeoff and landing.
  • the forward engines may provide horizontal thrust, while the propellers of the aft engines may be stowed at a fixed position in order to minimize drag.
  • the aft engines may be actively stowed with position monitoring.
  • Transition from vertical flight to horizontal flight and vice-versa may be accomplished via the tilt propeller subsystem.
  • the tilt propeller subsystem may redirect thrust between a primarily vertical direction during vertical flight phase (e.g., hover-phase) to a horizontal or near-horizontal direction during a forward-flight cruising phase, based on a tilt of one or more propellers (e.g., determining directionality of one or more propellers).
  • a variable pitch mechanism may change the forward engine’s propeller-hub assembly blade collective angles for operation during phases of flight, such as a hover-phase, transition phase, and cruisephase.
  • Vertical lift may be thrust in a primarily vertical direction (e.g., during a hover-phase).
  • Horizontal thrust may be thrust in a primarily horizontal direction (e.g., during a cruise-phase).
  • a “phase of flight,” or “flight mode,” e.g., hover, cruise, forward flight, takeoff, landing, transition
  • a combination flight conditions e.g., a combination of flight conditions within particular ranges
  • the forward engines may provide horizontal thrust for wing-borne take-off, cruise, and landing, and the wings may provide vertical lift.
  • the aft engines may not be used for generating thrust during a CTOL mission and the aft propellers may be stowed in place. In other embodiments, the aft engines may be used at reduced power to shorten the length of the CTOL takeoff or landing.
  • embodiments of the aircraft may include electrical systems responsible for power delivery and allocation.
  • electrical systems responsible for power delivery and allocation In an electric VTOL aircraft, the precise allocation of power to components of the high-voltage bus is of paramount importance for several critical reasons. Foremost among these is safety: ensuring that essential systems e.g. flight controls, avionics, and propulsion are consistently supplied with adequate power is vital to preventing operational failures that could lead to catastrophic consequences. Additionally, efficient power distribution is essential for minimizing energy losses and maximizing aircraft performance, which is particularly crucial given the battery-dependent nature of an electric VTOL aircraft, where range and operational capabilities are directly impacted by energy efficiency.
  • proper power allocation aids in effective thermal management by preventing the overheating of components and ensuring that cooling systems can maintain optimal temperatures. This not only enhances the reliability and longevity of the aircraft's systems but also supports the overall redundancy and reliability of the aircraft by ensuring that backup systems have sufficient power in the event of primary system failures. Moreover, strategic power distribution optimizes the weight and space of electrical systems, which is essential for maintaining the aircraft's efficiency given the strict weight limitations.
  • an aircraft of any of the disclosed embodiments may be simulated.
  • the aircraft may be in a simulated environment in a simulator (e.g., a simulator for flight training) or a virtual environment in a video game.
  • a display of an aircraft may be simulated.
  • the display e.g., control margin display
  • the display may be in a simulated environment in a simulator (e.g., a simulator for flight training) or a virtual environment in a video game.
  • a representation of the simulated display may be displayed on a display device (e.g., monitor, tablet, smartphone, computer screen, or any other display device) operatively connected to a processor configured to execute software code stored in a storage medium for performing flight controls operations, such as those further detailed below with reference to Fig. 10.
  • a display device e.g., monitor, tablet, smartphone, computer screen, or any other display device
  • a processor configured to execute software code stored in a storage medium for performing flight controls operations, such as those further detailed below with reference to Fig. 10.
  • FIG. 1 is an illustration of a perspective view of an exemplary VTOL aircraft, consistent with disclosed embodiments.
  • Fig- 2 is another illustration of a perspective view of an exemplary VTOL aircraft in an alternative configuration, consistent with embodiments of the present disclosure.
  • Figs. 1 and 2 illustrate a VTOL aircraft 100, 200 in a cruise configuration and a vertical take-off, landing and hover configuration (also referred to herein as a “lift” configuration), respectively, consistent with embodiments of the present disclosure.
  • Elements corresponding to Figs. 1 and 2 may possess like numerals and refer to similar elements of the aircrafts 100, 200.
  • the aircraft 100, 200 may include a fuselage 102, 202, wings 104, 204 mounted to the fuselage 102, 202 and one or more rear stabilizers 106, 206 mounted to the rear of the fuselage 102, 202.
  • a plurality of lift propellers 112, 212 may be mounted to wings 104, 204 and may be configured to provide lift for vertical take-off, landing and hover.
  • a plurality of tilt propellers 114, 214 may be mounted to wings 104, 204 and may be tiltable (e.g., configured to tilt or alter orientation) between the lift configuration in which they provide a portion of the lift required for vertical take-off, landing and hovering, as shown in Fig.
  • tilt propeller lift configuration refers to any tilt propeller orientation in which the tilt propeller thrust is providing primarily lift to the aircraft
  • tilt propeller cruise configuration refers to any tilt propeller orientation in which the tilt propeller thrust is providing primarily forward thrust to the aircraft.
  • lift propellers 112, 212 may be configured for providing lift only, with all horizontal propulsion being provided by the tilt propellers.
  • lift propellers 112, 212 may be configured with fixed positions and may only generate thrust during take-off, landing and hover phases of flight.
  • tilt propellers 114, 214 may be tilted upward into a lift configuration in which thrust from propellers 114, 214 is directed downward to provide additional lift.
  • tilt propellers 114, 214 may tilt from their lift configurations to their cruise configurations.
  • the orientation of tilt propellers 114, 214 may be varied from an orientation in which the tilt propeller thrust is directed downward (to provide lift during vertical take-off, landing and hover) to an orientation in which the tilt propeller thrust is directed rearward (to provide forward thrust to aircraft 100, 200).
  • the tilt propellers assembly for a particular electric engine may tilt about an axis of rotation defined by a mounting point connecting the boom and the electric engine.
  • lift may be provided entirely by wings 104, 204.
  • lift propellers 112, 212 may be shut off.
  • the blades 120, 220 of lift propellers 112, 212 may be held in low-drag positions for aircraft cruising.
  • lift propellers 112, 212 may each have two blades 120, 220 that may be locked, for example while the aircraft is cruising, in minimum drag positions in which one blade is directly in front of the other blade as illustrated in Fig. 1. In some embodiments, lift propellers 112, 212 have more than two blades. In some embodiments, tilt propellers 114, 214 may include more blades 116, 216 than lift propellers 112, 212. For example, as illustrated in Figs. 1 and 2, lift propellers 112, 212 may each include, e.g., two blades, whereas and tilt propellers 114, 214 may each include more blades, such as the five blades shown. In some embodiments, each of the tilt propellers 114, 214 may have 2 to 5 blades, and possibly more depending on the design considerations and requirements of the aircraft.
  • the aircraft may include a single wing 104, 204 on each side of fuselage 102, 202 (or a single wing that extends across the entire aircraft). At least a portion of lift propellers 112, 212 may be located rearward of wings 104, 204 (e.g., rotation point of propeller is behind a wing from a bird’ s eye view) and at least a portion of tilt propellers 114, 214 may be located forward of wings 104, 204 (e.g., rotation point of propeller is in front of a wing from a bird’s eye view).
  • lift propellers 112, 212 may be located rearward of wings 104, 204 (e.g., rotation point of propeller is behind a wing from a bird’ s eye view) and at least a portion of tilt propellers 114, 214 may be located forward of wings 104, 204 (e.g., rotation point of propeller is in front of a wing from a bird’s eye view).
  • all of lift propellers 112, 212 may be located rearward of wings 104, 204 and all of tilt propellers 114, 214 may be located forward of wings 104, 204. According to some embodiments, all lift propellers 112, 212 and tilt propellers 114, 214 may be mounted to the wings — e.g., no lift propellers or tilt propellers may be mounted to the fuselage. In some embodiments, lift propellers 112, 212 may be all located rearwardly of wings 104, 204 and tilt propellers 114, 214 may be all located forward of wings 104, 204. According to some embodiments, all lift propellers 112, 212 and tilt propellers 114, 214 may be positioned inwardly of the ends of the wing 104, 204.
  • lift propellers 112, 212 and tilt propellers 114, 214 may be mounted to wings 104, 204 by booms 122, 222.
  • Booms 122, 222 may be mounted beneath wings 104, 204, on top of the wings, and/or may be integrated into the wing profile.
  • lift propellers 112, 212 and tilt propellers 114, 214 may be mounted directly to wings 104, 204.
  • one lift propeller 112, 212 and one tilt propeller 114, 214 may be mounted to each boom 122, 222.
  • Lift propeller 112, 212 may be mounted at a rear end of boom 122, 222 and tilt propeller 114, 214 may be mounted at a front end of boom 122, 222. In some embodiments, lift propeller 112, 212 may be mounted in a fixed position on boom 122, 222. In some embodiments, tilt propeller 114, 214 may mounted to a front end of boom 122, 222 via a hinge. Tilt propeller 114, 214 may be mounted to boom 122, 222 such that tilt propeller 114, 214 is aligned with the body of boom 122, 222 when in its cruise configuration, forming a continuous extension of the front end of boom 122, 222 that minimizes drag for forward flight.
  • aircraft 100, 200 may include, e.g., one wing on each side of fuselage 102, 202 or a single wing that extends across the aircraft.
  • the at least one wing 104, 204 is a high wing mounted to an upper side of fuselage 102, 202.
  • the wings include control surfaces, such as flaps, ailerons, and/or flaperons (e.g., configured to perform functions of both flaps and ailerons).
  • wings 104, 204 may have a profile that reduces drag during forward flight.
  • the wing tip profile may be curved and/or tapered to minimize drag.
  • rear stabilizers 106, 206 include control surfaces, such as one or more rudders, one or more elevators, and/or one or more combined rudder-elevators.
  • the wing(s) may have any suitable design for providing lift, directionality, stability, and/or any other characteristic beneficial for aircraft.
  • the wings have a tapering leading edge.
  • lift propellers 112, 212 or tilt propellers 114, 214 may be canted relative to at least one other lift propeller 112, 212 or tilt propeller 114, 214, where canting refers to a relative orientation of the rotational axis of the lift propeller/tilt propeller about a line that is parallel to the forward-rearward direction, analogous to the roll degree of freedom of the aircraft.
  • one or more lift propellers 112, 212 and/or tilt propellers 114, 214 may canted relative to a cabin of the aircraft, such that the rotational axis of the propeller in a lift configuration is angled away from an axis perpendicular to the top surface of the aircraft.
  • the aircraft is a flying wing aircraft as shown in Fig. 9E below, and some or all of the propellers are canted away from the cabin.
  • Fig. 3 is an illustration of a top plan view of an exemplary VTOL aircraft, consistent with embodiments of the present disclosure. Aircraft 300 shown in the figure may be a top plan view of the aircraft 100, 200 shown in Figs. 1 and 2, respectively.
  • an aircraft 300 may include twelve electric propulsion systems distributed across the aircraft 300.
  • a distribution of electric propulsion systems may include six forward electric propulsion systems 314 and six aft electric propulsion systems 312 mounted on booms 322 forward and aft of the main wings 304 of the aircraft 300.
  • a length of the rear end of the boom 324 from the wing 304 to a lift propeller (part of electric propulsion system 312) may comprise a similar rear end of the boom 324 length across the numerous rear ends of the booms.
  • the length of the rear ends of the booms may vary, for example, across the six rear ends of the booms. Further, Fig.
  • FIG. 3 depicts an exemplary embodiment of a VTOL aircraft 300 with forward propellers (part of electric propulsion system 314) in a horizontal orientation for horizontal flight and aft propeller blades 320 in a stowed position for a forward phase of flight.
  • FIG. 4 is a schematic diagram illustrating exemplary propeller rotation of a VTOL aircraft, consistent with disclosed embodiments.
  • Aircraft 400 shown in the figure may be a top plan view of the aircraft 100, 200, and 300 shown in Figs. 1, 2, and 3, respectively.
  • An aircraft 400 may include six forward electric propulsion systems mounted on booms 422 with three of the forward electric propulsion systems being of CW type 424 and the remaining three forward electric propulsion systems being of CCW type.
  • three aft electric propulsion systems may be of CCW type 428 with the remaining three aft electric propulsion systems being of CW type 430.
  • Some embodiments may include an aircraft 400 possessing four forward electric propulsion systems and four aft electric propulsion systems, each with two CW types and two CCW types.
  • propellers may counterrotate with respect to adjacent propellers to cancel torque steer, generated by the rotation of the propellers, experienced by the fuselage or wings of the aircraft.
  • the difference in rotation direction may be achieved using the direction of engine rotation.
  • the engines may all rotate in the same direction, and gearing may be used to achieve different propeller rotation directions.
  • Fig. 5 is a schematic diagram illustrating exemplary power connections in a VTOL aircraft, consistent with disclosed embodiments.
  • a VTOL aircraft may have multiple power systems connected to diagonally opposing electric propulsion systems.
  • the power systems may include high voltage power systems.
  • Some embodiments may include high voltage power systems connected to electric engines via high voltage channels.
  • an aircraft 500 may include six power systems (e.g., battery packs), including power systems 526, 528, 530, 532, 534, and 536 stored within the wing 570 of the aircraft 500.
  • the power systems may power electric propulsion systems and/or other electric components of the aircraft 500.
  • the aircraft 500 may include six forward electric propulsion systems having six electric engines 502, 504, 506, 508, 510, and 512 and six aft electric propulsion systems having six electric engines 514, 516, 518, 520, 522, and 524.
  • the forward electric engines 502, 504, 506, 508, 510, and 512 may be connected to including power systems 526, 528, 530, 532, 534, and 536 via connections 538, 544, 550, 556, 560 and 564.
  • the aft electric engines 514, 516, 518, 520, 522, and 524 may be connected to including power systems 526, 528, 530, 532, 534, and 536 via connections 540, 546, 552, 558, 562 and 566.
  • power systems 526, 528, 530, 532, 534, and 536 may be interconnected via connections 542, 548 and 554.
  • one or more power systems may include a battery management system (“BMS”) (e.g., one BMS for each battery pack). While six power systems are shown in Fig. 5, the aircraft 500 may include any number and/or configuration of power systems.
  • BMS battery management system
  • the one or more battery management systems may communicate with a Flight Control System (“FCS”) of the aircraft (e.g., FCS 612 shown in Fig 6).
  • FCS Flight Control System
  • the FCS may monitor the status of one or more battery packs and/or provide commands to the one or more battery management systems which make corresponding adjustments to the high voltage power supply.
  • Fig. 6 illustrates block diagram of an exemplary architecture and design of an electric propulsion unit 600 consistent with disclosed embodiments.
  • Exemplary electric propulsion unit 600 includes an electric propulsion system 602, which may be configured to control aircraft propellers.
  • Electric propulsion system 602 may include an electric engine subsystem 604 that may supply torque, via a shaft, to a propeller subsystem 606 to produce the thrust of the electric propulsion system 602.
  • Some embodiments may include the electric engine subsystem 604 receiving low voltage direct current (LV DC) power from a Low Voltage System (LVS) 608.
  • LV DC low voltage direct current
  • LVS Low Voltage System
  • the electric engine subsystem 604 may be configured to receive high voltage (HV) power from a High Voltage Power System (HVPS) 610 comprising at least one battery or other device capable of storing energy.
  • HV power may refer to power that is lower in voltage than voltage provided by Low Voltage System (LVS) 608.
  • Some embodiments may include an electric propulsion system 602 including an electric engine subsystem 604 receiving signals from and sending signals to a flight control system 612.
  • a flight control system (FCS) 612 may comprise a flight control computer (FCC) capable of using Controller Area Network (“CAN”) data bus signals to send commands to the electric engine subsystem 604 and receive status and data from the electric engine subsystem 604.
  • FCC may include a device configured to perform one or more operations (e.g., computational operations) for an aircraft, such as at least one processor and a memory component, which may store instructions executable by the at least one processor to perform the operations, consistent with disclosed embodiments.
  • CAN data bus signals are used between the flight control computer and the electric engine(s), some embodiments may include any form of communication with the ability to send and receive data from a flight control computer to an electric engine. Some embodiments may include electric engine subsystems 604 capable of receiving operating parameters from and communicating operating parameters to an FCC in FCS 612, including speed, voltage, current, torque, temperature, vibration, propeller position, and/or any other value of operating parameters.
  • a 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 propellers.
  • TPS Tilt Propeller System
  • a tilt propeller system (TPS) 614 may include an apparatus capable of communicating operating parameters to an electric engine subsystem 604 and articulating an orientation of the propeller subsystem 606 to redirect the thrust of the tilt propellers during various phases of flight using mechanical means such as a gearbox assembly, linear actuators, and any other configuration of components to alter an orientation of the propeller subsystem 606.
  • electric engine subsystem may communicate an orientation of the propeller system (e.g., an angle between lift and forward thrust) to TPS 614 and/or FCS 612 (e.g., during flight).
  • a flight control system may include a system capable of controlling control surfaces and their associated actuators in an exemplary VTOL aircraft.
  • Fig- 7 is an illustration of a top plan view of an exemplary VTOL aircraft, consistent with embodiments of the present disclosure.
  • Aircraft 700 shown in the figure may be a top plan view of the aircraft 100, 200 shown in Figs. 1 and 2, respectively.
  • the control surfaces may include, in addition to the propeller blades discussed earlier, flaperons 712 and ruddervators 714.
  • Flaperons 712 may combine functions of one or more flaps, one or more ailerons, and/or one or more spoilers.
  • Ruddervators 714 may combine functions or one or more rudders and/or one or more elevators. Additionally or alternatively, control surfaces may include separate rudders and elevators. In aircraft 700, the actuators may include, in addition to the electric propulsion systems discussed earlier, control surface actuators (CSAs) associated with flaperons 712 and ruddervators 714, as discussed further below with reference to Fig. 8. [0072] Fig. 8 illustrates a flight control signaling architecture for controlling the control surfaces and associated actuators, according to various embodiments. Although Fig.
  • control surfaces and actuators may be controlled by a combination of four flight control computers (FCCs) — Left FCC, Lane A (L FCC-A), Left FCC, Lane B (L FCC-B), Right FCC, Lane A (R FCC-A), and Right FCC, Lane A (R FCC-B), although any other suitable number of FCCs may be utilized.
  • FCCs flight control computers
  • the FCCs may each individually control all control surfaces and actuators or may do so in any combination with each other.
  • each FCC may include one or more hardware computing processors.
  • each FCC may utilize a single-threaded computing process or a multi -threaded computing process to perform the computations required to control the control surfaces and actuators.
  • all computing process required to control the control surfaces and actuators may be performed on a single computing thread by a single flight control computer.
  • the FCCs may provide control signals to the control surface actuators, including the EPU inverters, TPACs, BMSs, flaperon CSAs, and ruddervator CSAs, via one or more bus systems.
  • the FCC may provide control signals, such as voltage or current control signals, and control information may be encoded in the control signals in binary, digital, or analog form.
  • the bus systems may each be a CAN bus system, e.g., Left CAN bus 1, Left CAN bus 2, Right CAN bus 1, Right CAN bus 2, Center CAN bus 1, Center CAN bus 2 (see Fig. 8).
  • multiple FCCs may be configured to provide control signals via each CAN bus system, and each FCC may be configured to provide control signals via multiple CAN bus systems.
  • L FCC-A may provide control signals via Left CAN bus 1 and Right CAN bus 1
  • L FCC-B may provide control signals via Left CAN bus 1 and Center CAN bus 1
  • R FCC-A may provide control signals via Center CAN bus 2 and Right CAN bus 2
  • R FCC-B may provide control signals via Left CAN bus 2 and Right CAN bus 2.
  • Figs. 9A-9F are illustrations of a top plan view of exemplary VTOL aircrafts, consistent with embodiments of the present disclosure.
  • design considerations cost, weight, size, performance capability etc.
  • power allocation to propellers may affect how forces are created. Therefore, the flight control system may adjust power delivery to various components in certain ways (e.g., those discussed in disclosed embodiments) to ensure safe and efficient operation of the aircraft.
  • Fig. 9A illustrates an arrangement of electric propulsion units, consistent with embodiments of the present disclosure.
  • 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 twelve electric propulsion systems distributed across the aircraft.
  • a distribution of electric propulsion systems may include six forward electric propulsion systems (901, 902, 903, 904, 905, and 906) and six aft electric propulsion systems (907, 908, 909, 910, 911, and 912).
  • the six forward electric propulsion systems may be operatively connected to tilt propellers and the six aft electric propulsion systems may be operatively connected to lift propellers.
  • the six forward electric propulsion systems and a number of aft electric propulsion systems may be operatively connected to tilt propellers and the remaining aft electric propulsion systems may be operatively connected to lift propellers. In other embodiments, all forward and aft electric propulsion systems may be operatively coupled to tilt propellers.
  • Fig. 9B illustrates an alternate arrangement of electric propulsion units, consistent with embodiments of the present disclosure.
  • 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 eight electric propulsion systems distributed across the aircraft.
  • a distribution of electric propulsion systems may include four forward electric propulsion systems (913, 914, 915, and 916) and four aft electric propulsion systems (917, 918, 919, and 920).
  • the four forward electric propulsion systems may be operatively connected to tilt propellers and the four aft electric propulsion systems may be operatively connected to lift propellers.
  • the four forward electric propulsion systems and a number of aft electric propulsion systems may be operatively connected to tilt propellers and the remaining aft electric propulsion systems may be operatively connected to lift propellers. In other embodiments, all forward and aft electric propulsion systems may be operatively coupled to tilt propellers.
  • Fig. 9C illustrates an alternate arrangement of electric propulsion units, consistent with embodiments of the present disclosure.
  • 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.
  • a distribution of electric propulsion systems may include a first set of four electric propulsion systems 921, 922, 923, and 924 coplanar in a first plane and a second set of two electric propulsion systems 925 and 926 coplanar in a second plane.
  • the first set of electric propulsion systems 921-924 may be operatively connected to tilt propellers and second set of electric propulsion systems 925 and 926 may be operatively connected to lift propellers. In other embodiments, the first set of electric propulsion systems 921-924 and the second set of aft electric propulsion systems 925 and 926 may all be operatively connected to tilt propellers.
  • Fig. 9D illustrates an alternate arrangement of electric propulsion units, consistent with embodiments of the present disclosure.
  • 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.
  • a distribution of electric propulsion systems may include four coplanar electric propulsion systems 927, 928, 929, and 930.
  • all of the electric propulsion systems may be operatively connected to tilt propellers.
  • Fig. 9E illustrates an alternate arrangement of electric propulsion units, consistent with embodiments of the present disclosure.
  • 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.
  • the aircraft may include four forward electric propulsion systems 931, 932, 933, and 934 operatively connected to tilt propellers and the two aft ducted fans 935 and 936 operatively connected to lift propellers.
  • the aircraft may include ten electric propulsion systems distributed across the aircraft.
  • the aircraft may include six forward electric propulsion systems operatively connected to tilt propellers and the four aft electric propulsion systems operatively connected to lift propellers. In some embodiments, some or all of the aft electric propulsion systems may operatively connected to tilt propellers.
  • the aircraft may have a flying wing configuration, such as a tailless fixed-wing aircraft with no definite fuselage. In some embodiments, the aircraft may have a flying wing configuration with the fuselage integrated into the wing.
  • the tilt propellers may rotate in a plane above the body of the aircraft when the tilt propellers operate in a lift configuration.
  • Fig. 9F illustrates an alternate arrangement of electric propulsion units, consistent with the embodiments of the present disclosure.
  • the aircraft may be atop plan view of an exemplary aircraft.
  • the aircraft may include ducted fans 936, 937, 938, and 939 operably connected to the electric propulsion systems.
  • the aircraft may include a bank of ducted fans on each wing of the aircraft and the bank of ducted fans may be connected to tilt together (e.g., between lift and forward thrust configuration).
  • the aircraft includes a left and right front wing and a left and right rear wing.
  • each wing of the aircraft includes a bank of connected ducted fans.
  • each bank of connected ducted fans are tiltable (e.g., between lift and forward thrust), while in other embodiments only the bank of fans on the front wing(s) are tiltable.
  • the forward electric propulsion systems and aft electric propulsion systems may be of a clockwise (CW) type or counterclockwise (CCW) type. Some embodiments may include various forward electric propulsion systems possessing a mixture of both CW and CCW types. In some embodiments, the aft electric propulsion systems may possess a mixture of CW and CCW type systems among the aft electric propulsion systems. In some embodiments, each electric propulsion systems may be fixed as clockwise (CW) type or counterclockwise (CCW) type, while in other embodiments, one or more electric propulsion systems may vary between clockwise (CW) and counterclockwise (CCW) rotation.
  • Fig. 10 illustrates a functional block diagram of an exemplary control system 1000 of an aircraft, consistent with disclosed embodiments.
  • System 1000 may be implemented by at least one processor (e.g., at least one a 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 a flight control system of the aircraft (e.g., part of FCS 612 in Fig. 6) and may be configured to perform a single step or sequence repeatedly until a desired or commanded outcome is achieved.
  • System 1000 further includes one or more storage mediums storing model(s), function(s), table(s), and/or any information for executing the disclosed processes.
  • any or each box indicating a command model e.g., 1004, 1006, 1008, and 1010
  • feedback 1012, 1016, 1018, and 1022
  • feed forward 1014, 1020
  • Outer Loop Allocation 1024, 1026
  • module(s), script(s), function(s), application(s), and/or program(s) that are executed by processor(s) and/or microprocessor(s) of system 1000. It is appreciated that the complexity and interconnectedness of the functional block diagram of Fig. 10 would be impossible, or at least impractical, to effectively implement by a human user, especially when considering that these functionalities are implemented while the aircraft is flying (including taking off or landing).
  • control system 1000 may be configured based on one or more flight control laws.
  • Flight control law may comprise a set of algorithms, models, and/or rules configured to govern a behavior of an aircraft (e.g., control or influence one or more effectors of the aircraft) in response to one or more pilot inputs and external factors.
  • flight control laws may be configured to achieve at least one of desired flight characteristics, stability, or performance.
  • flight control laws may be configured to ensure stability and controllability of an aircraft by controlling how the aircraft responds to at least one of one or more pilot inputs, vehicle dynamics (e.g., disturbances, such as turbulence, gusts, etc.), or changes in flight conditions (e.g., altitude, airspeed, angle of attack).
  • System 1000 may detect one or more inputs, such as from a pilot input device configured to receive at least one pilot input and generate or influence a signal.
  • a pilot input may be generated by and/or received from an input device or mechanism of the aircraft, such as a button, a switch, a stick, a slider, an inceptor, or any other device configured to generate or influence a signal based on a physical action from a pilot.
  • a pilot input device may include one or more of right inceptor(s) (e.g., moving right inceptor left/right 1002a and/or right inceptor forward/aft 1002e), left inceptor(s) (e.g., moving left inceptor left/right 1002c and/or left inceptor forward/aft 1002g), and/or left inceptor switch 1002f.
  • a pilot input device may include an interface with an autopilot system (e.g., display screen(s), switch(es), button(s), lever(s), and/or other interface(s)).
  • system 1000 may further detect inputs from an autopilot system, such as autopilot roll command 1002b, autopilot climb command 1002d, and/or other command(s) to control the aircraft.
  • the one or more inputs may include at least one of a position and/or rate of a right inceptor and/or a left inceptor, signals received (e.g., response type change commands, trim inputs, reference inputs, backup control inputs, etc.) from switches on the inceptors, measurements of aircraft state and environmental conditions (e.g., measured load factor, airspeed, roll angle, pitch angle, actuator states, battery states, aerodynamic parameters, temperature, gusts, etc.) based on data received from one or more sensors of the aircraft, obstacles (e.g., presence or absence of other aircraft and/or debris), and an aircraft mode (e.g., taxiing on the ground, takeoff, in-air).
  • signals received e.g., response type change commands, trim inputs, reference inputs, backup control inputs, etc.
  • measurements of aircraft state and environmental conditions e.g., measured load factor, airspeed, roll angle, pitch angle, actuator states, battery states, aerodynamic parameters, temperature, gusts, etc.
  • obstacles e.g.
  • right inceptor L/R 1002a may comprise a lateral position and/or rate of a right inceptor (e.g., an inceptor positioned to the right of another inceptor and/or an inceptor positioned on the right side of a pilot area)
  • autopilot roll command 1002b may comprise a roll signal received in autopilot mode
  • left inceptor L/R 1002c may comprise a lateral position and/or rate of a left inceptor (e.g., an inceptor positioned to the left of another inceptor and/or an inceptor positioned on the left side of a pilot area)
  • autopilot climb command 1002d may comprise a climb signal received in autopilot mode
  • right inceptor F/A 1002e may comprise a longitudinal position and/or rate of the right inceptor
  • left inceptor switch 1002f may comprise a signal from a switch for enabling or disabling automatic transition function 1003
  • left inceptor F/A 1002g may comprise a longitudinal position and
  • Each input may include data as listed above (e.g., signals from switches, measurements of aircraft state, aircraft mode, etc.).
  • Actuator states may include actuator hardware limits, such as travel limits, speed limits, response time limits, etc., and can include actuator health indicators that may indicate deteriorations in actuator performance that may limit a given actuator’s ability to satisfy actuator commands. Actuator states may be used to determine the bounds (e.g., minimum/maximum values) for individual actuator commands.
  • Battery states may correspond to remaining energy of the battery packs of the aircraft, which may be monitored when control allocation 1029 considers balancing battery pack energy states.
  • Aerodynamic parameters may be parameters derived from aerodynamic and acoustic modeling and can be based on the actuator Jacobian matrices and actuator states.
  • Each input received from an inceptor may indicate a corresponding adjustment to an aircraft’s heading or power output.
  • Command models 1004, 1006, 1008 and 1010 may be configured to determine a shape (e.g., aggressiveness, slew rate, damping, overshoot, etc.) of an ideal aircraft response.
  • each command model of command models 1004, 1006, 1008 and 1010 may be configured to receive and interpret at least one of inputs 1002a, 1002b, 1002c, 1002d, 1002e, 1002f and 1002g and, in response, compute a corresponding change to an aircraft’s orientation, heading, and propulsion, or a combination thereof using an integrator (not pictured).
  • right inceptor L/R 1002a and autopilot roll command 1002b may be fed into turn -rate command model 1004, left inceptor L/R 1002c may be fed into lateral speed command model 1006, autopilot climb command 1002d and right inceptor F/A 1002e may be fed into climb command model 1008, and left inceptor F/A 1002g may be fed into forward speed command model 1010.
  • an output from automatic transition function 1003 may be fed into at least one of climb command model 1008 or forward speed command model 1010. For example, based on receiving an enable signal from left inceptor switch 1002f, automatic transition function 1003 may automatically determine at least one of a climb signal or a forward speed signal for transmission to at least one of climb command model 1008 or forward speed command model 1010.
  • Turn-rate command model 1004 may be configured to output a desired position and/or turn-rate command and may also be configured to compute a desired heading of the aircraft to be assumed when the inceptor is brought back to a centered position (e.g., in detent).
  • Lateral speed command model 1006 may be configured to output a desired position and/or lateral speed command.
  • Climb command model 1008 may be configured to output at least one of a desired altitude, vertical speed, or vertical acceleration command.
  • Forward speed command model 1010 may be configured to output at least one of a desired position, longitudinal speed, or longitudinal acceleration command.
  • one or more of the command models may be configured to output an acceleration generated in response to changes in speed command.
  • climb command model 1008 may be configured to output a vertical acceleration generated in response to a change in vertical speed command.
  • Feed forward 1014 and 1020 may each receive as input one or more desired changes (e.g., desired position, speed and/or acceleration) from corresponding command models 1004, 1006, 1008 or 1010 as well as data received from the one or more aircraft sensors (e.g., airspeed, vehicle orientation, vehicle load factor, measured acceleration, vehicle mass and inertia, air density, altitude, aircraft mode, etc.) and may be configured to output, for each desired change, a corresponding force to accomplish the desired change.
  • feed forward 1014 and 1020 may be configured to determine the corresponding force using simplified models of aircraft dynamics.
  • feed forward 1014 and 1020 may be configured to determine a force to cause the aircraft to follow a desired acceleration command.
  • feed forward 1014 and 1020 may be configured to use a model predicting an amount of drag on the vehicle produced as a function of speed in order to determine a force required to follow a desired speed command signal.
  • Feedback 1012, 1016, 1018, and 1022 may each receive as input the one or more desired changes (e.g., desired position, speed and/or acceleration) from command models 1004, 1006, 1008 and 1010 as well as data received from Vehicle Sensing 1031 indicative of Vehicle Dynamics 1030.
  • sensed Vehicle Dynamics 1030 may comprise the physics and/or natural dynamics of the aircraft, and Vehicle Sensing 1031 sensor measurements may capture how the aircraft moves in response to pilot inputs, propulsion system outputs or ambient conditions.
  • Vehicle Dynamics 1030 may represent the control of different flight elements (e.g., electric propulsion system(s) and/or control surfaces) and the corresponding effect on the flight elements and aircraft dynamics.
  • data received from Vehicle Sensing 1031 may include error signals generated, by one or more processors, based on exogenous disturbances (e.g., gust causing speed disturbance).
  • feedback 1012, 1016, 1018 and 1022 may be configured to generate feedback forces (e.g., at an actuator) based on the received error signals.
  • feedback 1012, 1016, 1018 and 1022 may generate feedback forces with the intent of counteracting the effect(s) of external disturbances. Additionally or alternatively, feedback 1012, 1016, 1018 and 1022 may be configured to generate feedback forces based on modeling errors. For example, if an incorrect aircraft mass is input into either feed forward 1014 or 1020, the aircraft may accelerate faster or slower than the desired change. Based on determining a difference between the desired acceleration and the measured acceleration, one or more processors may generate an error signal (e.g., included in Vehicle Sensing 1031) which may be looped into feedback 1012, 1016, 1018 or 1022 to determine an additional force needed to correct the error.
  • an error signal e.g., included in Vehicle Sensing 1031
  • feedback 1012, 1016, 1018 or 1022 may be disabled.
  • system 1000 in response to losing position and/or ground speed feedback due to disruption of global position system (GPS) communication, system 1000 may be configured to operate without feedback 1012, 1016, 1018 or 1022 until GPS communication is reconnected.
  • GPS global position system
  • feedback 1012, 1016, 1018 or 1022 may receive as input a plurality of measurements as well as a trust value for each measurement indicating whether the measurement is valid.
  • one or more processors of system 1000 may assign a Boolean (true/false) value for each measurement used in system 1000 to indicate that the measurement is trustworthy (e.g., yes) or that the measurement may be invalid (e.g., no). Based on one or more processors identifying a measurement as invalid, feedback 1012, 1016, 1018 or 1022 may omit that measurement for further processing.
  • feedback 1012, 1016, 1018 or 1022 may omit subsequent heading measurements in determining feedback force(s).
  • feedback 1012, 1016, 1018 or 1022 may determine one or more feedback forces based on actuator state information received from one or more sensors (e.g., included in Vehicle Sensing 1031).
  • actuator state information indicating that there is a failure of an actuator
  • one or more processors of system 1000 may update one or more processes of System 1000 and determine an alternative command to achieve the desired change.
  • one or more processors of system 1000 may adjust one or more model(s), function(s), algorithm(s), table(s), input(s), parameter(s), threshold(s), and/or constraint(s) in response to the failure of an actuator.
  • Alternative command(s) e.g., yaw, pitch, roll, thrust, or torque
  • one or more processors of system 1000 may update one or more processes of system 1000 (e.g., as described above) and determine an alternative command to achieve the desired change.
  • Total desired forces may be calculated based on outputs of feedback 1012, 1016, 1018 and 1022 and feed forward 1014 and 1020.
  • one or more processors of system 1000 may calculate a desired turn -rate force by summing the outputs of feedback 1012 and feed forward 1014. Additionally or alternatively, one or more processors of system 1000 may calculate a desired lateral force by summing the outputs of feedback 1016 and feed forward 1014. Additionally or alternatively, one or more processors of system 1000 may calculate a desired vertical force by summing the outputs of feedback 1018 and feed forward 1020. Additionally or alternatively, one or more processors of system 1000 may calculate a desired longitudinal force by summing the outputs of feedback 1022 and feed forward 1020.
  • Lateral/Directional Outer Loop Allocation 1024 and Longitudinal Outer Loop Allocation 1026 may each be configured to receive as input one or more desired forces and data received from Vehicle Sensing 1031 (e.g., airspeed, vehicle orientation, vehicle load factor, measured acceleration, vehicle mass and inertia, indications of working/failed actuators, air density, altitude, aircraft mode, whether the aircraft is in the air or on the ground, weight on wheels, etc.). Based on the inputs, Outer Loop Allocation 1024 and 1026 may be configured to command roll, command yaw, command pitch, demand thrust, or output a combination of different commands/demands in order to achieve the one or more desired forces.
  • Vehicle Sensing 1031 e.g., airspeed, vehicle orientation, vehicle load factor, measured acceleration, vehicle mass and inertia, indications of working/failed actuators, air density, altitude, aircraft mode, whether the aircraft is in the air or on the ground, weight on wheels, etc.
  • Lateral/Directional Outer Loop Allocation 1024 may receive as input a desired turn-rate force and/or a desired lateral force and may command roll or command yaw. In some embodiments, Lateral/Directional Outer Loop Allocation 1024 may determine output based on a determined flight mode. A flight mode may be determined using pilot inputs (e.g., a selected mode on an inceptor) and/or sensed aircraft information (e.g., an airspeed).
  • pilot inputs e.g., a selected mode on an inceptor
  • sensed aircraft information e.g., an airspeed
  • Lateral/Directional Outer Loop Allocation 1024 may determine a flight mode of the aircraft using at least one of a determined (e.g., sensed or measured) airspeed or an input received at a pilot inceptor button (e.g., an input instructing the aircraft to fly according to a particular flight mode).
  • a pilot inceptor button e.g., an input instructing the aircraft to fly according to a particular flight mode.
  • Lateral/Directional Outer Loop Allocation 1024 may be configured to prioritize a pilot inceptor button input over measured airspeed in determining the flight mode (e.g., the pilot inceptor button is associated with a stronger weight or higher priority than a measured airspeed).
  • Lateral/Directional Outer Loop Allocation 1024 may be configured to blend (e.g., using weighted summation) the determined airspeed and pilot inceptor button input to determine the flight mode of the aircraft. In a hover flight mode, Lateral/Directional Outer Loop Allocation 1024 may achieve the desired lateral force with a roll command (e.g., roll angle, roll rate) and may achieve the desired turn-rate force with a yaw command. In some embodiments, such as in hover flight mode, the aircraft may be configured to not be able to accelerate outside a predetermined hover envelope (e.g., hover speed range).
  • a predetermined hover envelope e.g., hover speed range
  • Lateral/Directional Outer Loop Allocation 1024 may achieve the desired lateral force with a yaw command and may achieve the desired turn-rate force with a roll command.
  • Lateral/Directional Outer Loop Allocation 1024 may be configured to determine output based on sensed airspeed.
  • Lateral/Directional Outer Loop Allocation 1024 may achieve desired forces using a combination of a roll command and a yaw command.
  • Longitudinal Outer Loop Allocation 1026 may receive as input a desired vertical force and/or a desired longitudinal force and may output at least one of a pitch command (e.g., pitch angle) or a thrust vector demand.
  • a thrust vector demand may include longitudinal thrust (e.g., mix of nacelle tilt and front propeller thrust) and vertical thrust (e.g., combined front and rear thrust).
  • Longitudinal Outer Loop Allocation 1026 may determine output based on a determined flight mode. For example, in a hover flight mode, Longitudinal Outer Loop Allocation 1026 may achieve a desired longitudinal force by lowering a pitch attitude and by using longitudinal thrust, and may achieve a desired vertical force with vertical thrust.
  • Longitudinal Outer Loop Allocation 1026 may achieve a desired longitudinal force with longitudinal thrust (e.g., front propeller thrust). In a cruise flight mode, Longitudinal Outer Loop Allocation 1026 may achieve a desired vertical force by commanding pitch (e.g., raising pitch attitude) and demanding thrust (e.g., increasing longitudinal thrust).
  • Inner loop control laws 1028 may be configured to determine moment commands based on at least one of a roll command, yaw command, or pitch command from Lateral/Directional Outer Loop Allocation 1024 or Longitudinal Outer Loop Allocation 1026.
  • Inner loop control laws 1028 may be dependent on sensed Vehicle Dynamics (e.g., from Vehicle Sensing 1031). For example, Inner loop control laws 1028 may be configured to compensate for disturbances at the attitude and rate level in order to stabilize the aircraft. Additionally or alternatively, Inner loop control laws 1028 may consider periods of natural modes (e.g., phugoid modes) that affect the pitch axis, and may control the aircraft appropriately to compensate for such natural modes of the vehicle. In some embodiments, inner loop control laws 1028 may be dependent on vehicle inertia.
  • Inner loop control laws 1028 may determine moment commands using one or more stored dynamics models that reflect the motion characteristics of the aircraft (e.g., the aerodynamic damping and/or inertia of the aircraft). In some embodiments, the Inner loop control laws 1028 may use a dynamic model (e.g., a low order equivalent system model) to capture the motion characteristics of the aircraft and determine one or more moments that will cause the aircraft to achieve the commanded roll, yaw, and/or pitch.
  • a dynamic model e.g., a low order equivalent system model
  • Some embodiments may include determining (e.g., by inner loop control laws 1028 or other component) a moment command based on at least one received command (e.g., a roll command, yaw command, and/or pitch command) and a determined (e.g., measured) aircraft state.
  • a moment command may be determined using a difference in the commanded aircraft state and the measured aircraft state.
  • a moment command may be determined using the difference between a commanded roll angle and a measured roll angle.
  • Control Allocation 1029 may control the aircraft (e.g., through flight elements) based on the determined moment command(s).
  • Control Allocation 1029 may control (e.g., transmit one or more commands to) one or more electric propulsion system(s) of the aircraft (e.g., electric propulsion system 602 shown in Fig. 6), including tilt actuator(s), electric engine(s), and/or propeller(s).
  • Control Allocation 1029 may further control one or more control surface(s) of the aircraft (e.g., control surfaces, such as flaperons 712 and ruddervators 714 shown in Fig. 7), including flaperon(s), ruddervator(s), aileron(s), spoiler(s), rudder(s), and/or elevator(s).
  • a pilot inceptor input may create roll, yaw, pitch, and/or thrust commands.
  • a right inceptor may control roll and pitch and a left inceptor and/or pedal(s) may control yaw and thrust.
  • Control Allocation 1029 may accept as inputs one or more of force and moment commands, data received from the one or more aircraft sensors, envelope protection limits, scheduling parameter, and optimizer parameters. Control Allocation 1029 may be configured to determine, based on the inputs, actuator commands by minimizing an objective function that includes one or more primary objectives, such as meeting commanded aircraft forces and moments, and one or more secondary, which can include minimizing acoustic noise and/or optimizing battery pack usage.
  • control allocation 1029 may be configured to compute the limits of individual actuator commands based on the actuator states and envelope protection limits.
  • Envelope protection limits may include one or more boundaries that the aircraft should operate within to ensure safe and stable flight.
  • envelope protection limits may be defined by one or more of speed, altitude, angle of attack, or load factor.
  • envelope protection limits may include one or more bending moments or load constraints.
  • control allocation 1029 may use envelope protection limits to automatically adjust one or more control surfaces or control settings. Doing so may prevent the aircraft from undesirable scenarios such as stalling or structural strain or failure.
  • Control allocation 1029 sends commands to one or more flight elements to control the aircraft.
  • the flight elements will move in accordance with the controlled command.
  • Various sensing systems and associated sensors as part of Vehicle Sensing 1031 may detect the movement of the flight elements and/or the dynamics of the aircraft and provide the information to Feedback 1012, 1016, 1018, 1022, Outer Loop allocation 1024 and 1026, Inner Loop Control laws 1028, and Control Allocation 1029 to be incorporated into flight control.
  • Vehicle Sensing 1031 may include one or more sensors to detect vehicle dynamics. For example, Vehicle Sensing 1031 may capture how the aircraft moves in response to pilot inputs, propulsion system outputs or ambient conditions.
  • Vehicle Sensing 1031 may detect an error in the aircraft’s response based on exogenous disturbances (e.g., gust causing speed disturbance). Further, Vehicle Sensing 1031 may include one or more sensors to detect propeller speed, such as a magnetic sensor (e.g., Hall effect or inductive sensor) or an optical sensor (e.g., a tachometer) configured to detect the rotor speed of the aircraft engine (and thereby the speed of the propeller). Vehicle sensing 1031 may include one or more sensors to detect a nacelle tilt angle (e.g., a propeller rotation axis angle between a lift configuration (e.g., Fig. 2) and forward thrust configuration (e.g., Fig. 1)).
  • a nacelle tilt angle e.g., a propeller rotation axis angle between a lift configuration (e.g., Fig. 2) and forward thrust configuration (e.g., Fig. 1)).
  • one or more magnetic sensors e.g., Hall effect or inductive sensor
  • position displacement sensors e.g., linear displacement sensors, and/or other sensor(s) associated with the tilt actuator
  • a tilt angle e.g., relative to the aircraft and/or wing
  • one or more pitot tubes, accelerometers, and/or gyroscopes may detect a pitch angle of the aircraft, which may be provided to system 1000.
  • Vehicle Sensing 1031 may combine tilt angle sensor measurements and aircraft pitch measurements to determine an overall nacelle tilt angle for the propellers.
  • Vehicle sensing 1031 may include one or more sensors configured to detect an engine torque and/or thrust, such as one or more current sensors or voltage sensors, strain gauges, load cells, and/or propeller vibration sensors (e.g., accelerometers).
  • Vehicle sensing 1031 may include one or more sensors configured to detect vehicle dynamics, such as acceleration and/or pitch orientation sensors (e.g., accelerometer(s), 3-axis accelerometer(s), gyroscope(s), 3-axis gyroscope(s), and/or tilt-position sensors to determine angles of engines) and airspeed sensors (e.g., pitot tube sensors). Vehicle sensing 1031 may further include one or more inertial measurement units (IMUs) to determine an aircraft state based on these measurements.
  • IMUs inertial measurement units
  • An aircraft state may refer to forces experienced by, an orientation of, a position of (e.g., altitude), and/or movement of, the aircraft.
  • an aircraft state may include at least one of: a position of the aircraft (e.g., a yaw angle, roll angle, pitch angle, and/or any other orientation across one or two axes), velocity of the aircraft, angular rate of the aircraft (e.g., roll, pitch, and/or yaw rate), and/or an acceleration of the aircraft (e.g., longitudinal, lateral and/or vertical acceleration), or any physical characteristic of the aircraft or one of its components.
  • a position of the aircraft e.g., a yaw angle, roll angle, pitch angle, and/or any other orientation across one or two axes
  • velocity of the aircraft e.g., angular rate of the aircraft (e.g., roll, pitch, and/or yaw rate)
  • an acceleration of the aircraft e.g., longitudinal, lateral and/or vertical acceleration
  • Vehicle Sensing 1031 may include an inertial navigation systems (INS) and/or an air data and/or an attitude heading reference systems (ADAHRS).
  • the inertial navigation systems (INS) and/or an air data and attitude heading reference systems (ADAHRS) may include one or more inertial measurement units (IMUs) and corresponding sensors (e.g., accelerometers, gyroscopes, three-axis gyroscopes, and/or three-axis accelerometers).
  • the INS and/or ADAHRS may filter and/or otherwise process sensor measurements to determine an aircraft state (e.g., acceleration or angular rate).
  • the INS and/or ADAHRS may determine angular rates based on gyroscope measurements and may determine acceleration based on measurements from an accelerometer.
  • a high-voltage (HV) bus may be a centralized electrical distribution network configured to distribute electrical power at high voltages.
  • a high-voltage bus may serve as a common connection point for multiple circuits, that may allow transfer of power between different parts of an electrical system.
  • the high-voltage bus may be connected to a plurality of battery cells, which may be responsible for providing power for at least some electrical systems of the aircraft.
  • HV bus may be comprised in high-voltage power system, for example HVPS 610 of Fig. 6.
  • the embodiments of the present disclosure may relate to a method for allocating power to component of HV bus.
  • the method for allocating power to components of a HV bus may be carried out by a HV protection function.
  • Some of the objectives of the HV protection function may be to control power flow between the HV bus and the electrical components it may be connected to in order to ensure safe operation of the aircraft, to prolong lifetime of power source(s) and components, and to ensure efficient operation of the aircraft.
  • HV protection function may control power delivered to components connected to the HV bus, as well as power received from the components of the HV bus. For example, in some flight scenario idle rotor and/or propellers may generate power that may be utilized for charging of batteries of the battery pack, which may also be overseen by the HV protection function.
  • Fig. 11 is a block diagram describing an exemplary computer method 1100 for allocating power to components of at least one high-voltage (HV) bus according to the embodiments of the present disclosure.
  • Method 1100 may be implemented using a particular machine (machine (e.g., an aircraft), according to various embodiments of the present disclosure, thereby improving the technology of the aircraft (e.g., aircraft safety, thermal efficiency, range capability, payload capacity, structural integrity).
  • the steps of method 1100 may be performable by, for example, system 1000 of Fig. 10 executing on or otherwise using the components of any aircraft of Figs. 1-8, 9A-9E, 10, or any flight control computer (e.g., a computer-implemented method) or flight control system.
  • a flight control computer of the aircraft may be configured to perform one or more steps of method 1100.
  • the complexity of method 1100 would be impossible, or at the very least grossly impractical, to effectively implement by a human user, especially when considering that these functionalities are implemented (e.g., in real time) while the aircraft is flying (including taking off or landing), while energy used by the aircraft and battery conditions are constantly changing.
  • power allocation e.g., using method 1100 improves the safety of the aircraft as well as lifespan of different components, without requiring any direct involvement of a user (e.g., pilot).
  • the steps of method 1100 may be activated or adjusted as needed based on aircraft flight conditions, maneuvers, or operational requirements. It is appreciated that the illustrated method 1100 can be altered to modify the order of steps and to include additional steps.
  • method 1100 may receive an input indicative of an operation command.
  • an operation command may be a pilot command (e.g., command requesting an aircraft and/or aircraft component action) or an input indicative of pilot command.
  • An operation command may relate to flight of the aircraft and may correspond to a command requesting execution of a maneuver by the aircraft. Realization of such operation command may require active involvement of effectors and/or actuators. For instance, execution of the requested maneuver may involve adjustment of thrust generated by at least some of the engines of the aircraft. Execution of such command may correspond to a request for power allocation to relevant components of the aircraft and this request may be analyzed to decide whether or how the requested power may be allocated to relevant components. When the power demand is too large (e.g., will exceed a safety constraint, temperature constraint, flight efficiency constraint, etc.), method 1100 may determine an alteration of power allocation with respect to the demand based on the received operation command.
  • method 1100 may calculate a total power demand associated with a plurality of components connected to the high-voltage bus based on the received input, where each component may have a respective component power demand.
  • calculating the total power demand may include adding respective component power demands, which may be determined according to calculations determined by system 1000 (e.g., according to one or more algorithmic operations discussed above). The calculation of the total power demand may be based on an efficiency of at least one (e.g., each) component. Efficiency may refer to how much input power is successfully converted to a desired output power. For example, efficiency of a component may be defined as the ratio of the useful mechanical power output (i.e power input to the engine minus energy losses, e.g.
  • efficiency may include a power ratio, such as a ratio of input power (e.g., power from a battery to an engine), which may be defined as current times voltage out of a battery, to output power (e.g., of an engine), which may be defined as an amount of torque produced by the engine times a revolutions per minute (RPM) of a rotor and/or propeller of the engine.
  • the electrical input power may include losses occurring between the battery output and the engine and/or propeller.
  • power demand of a component may be calculated by dividing a power demand of a component with no losses included by the actual efficiency of the component. As a result, calculated power demand may account for power losses associated with the component.
  • the efficiency of the component may be measured during operation of the component, while in some embodiments the efficiency of the component may be predicted based on a mathematical model of operation of the component.
  • the measurements of efficiency may be implemented by measuring power (e.g., using one or more sensors, such as a shunt) near the batteries (e.g., at contact points for the batteries) and at the intended delivery point (e.g. engine, rotor, and/or propeller, such as by using a measured torque and rpm value), which may enable accounting for power losses between the point of generation (e.g. batteries) and point of deliver (e.g. engine and/or rotor).
  • method 1100 may combine the efficiency measurements during operation of the component with the predictions based on a mathematical model of operation of the component, which may provide more accurate efficiency figures and to allow for redundancy in the system.
  • the plurality of components connected to the high- voltage bus may comprise at least one engine and the total power demand may be determined based on a torque, revolutions per minute (rpm) or power demand (or any combination thereof) of the at least one engine.
  • the HV protection function may determine that fulfilment of the command may require at least some of the engines connected to the HV bus to increase delivered torque.
  • an engine may power a rotor and/or propeller, which may rotate with a certain speed (e.g., RPM).
  • the power demand for such configuration may be calculated by multiplying toque required by the rotor and/or propeller by its rotational speed. As described above with respect to step 1120, the calculation may also include a factor representing efficiency of the engine and/or the rotor/propeller.
  • the HV protection function implementing method 1100 may calculate the amount of power required by each engine to realizes (e.g., execute, fulfil, perform) the operation command.
  • the total power demand may be based on power supplied by (e.g., generated by) at least one component that may also be configured to demand power, for example when an engine is regenerating power through a windmill effect of an associated propeller.
  • method 1100 may include determining the high-voltage bus power limit assigned to the high-voltage bus based on a hardware power limit and electrical power available from the plurality of battery cells.
  • the determination of the high-voltage bus power limit assigned to the high-voltage bus may be based on at least one of the measured state of charge, state of health or state of power of the battery cells, which may be measured by at least one of the battery management systems of Fig. 8.
  • determining the power limits of a battery pack may be based on parameters (e.g., monitored by the BMS) such as voltage, current, temperature, state of charge (SoC), state of health (SoH), state of power (SoP), and internal resistance to ensure voltage remains within manufacturer-specified thresholds and track voltage sag under load.
  • State of charge (SOC) may refer to an available battery pack capacity relative to the battery pack’s rated capacity.
  • the state of charge may be based on an open circuit voltage (OCV) of the battery pack, where the OCV is the resting voltage of the battery pack (e.g., based on a battery pack without current flow for a set period of time).
  • State of health may refer to an overall condition of a battery compared to its ideal conditions (e.g., amount of degradation).
  • method 1100 may include comparing (e.g., by the BMS) charging and discharging currents to recommended maximums and analyze historical current data for typical usage patterns, as well as it may also monitor cell and pack temperatures, setting limits to prevent overheating and thermal runaway.
  • State of power SOP
  • SoC SoC may be used by method 1100 to avoid overcharging and deep discharging, while SoH may provide information relating to battery aging and necessary power adjustments.
  • SOP may indicate the available constant discharge power over a projected time window Additionally, internal resistance may be tracked to manage power limits as the battery degrades.
  • step 1130 method 1100 may compare the total power demand to a high- voltage bus power limit assigned to the high-voltage bus to determine a difference between the high-voltage bus power limit and the total power demand.
  • Fig. 12 is a schematic diagram showing an exemplary scenario illustrating determination of the difference between the high- voltage bus power limit and the total power demand.
  • Graph 1210 shows power demand of four engines, where HV protection function implementing method 1100 may calculate total power demand associated with four engines connected to the HV bus based on the received input indicative of operation command. As shown in graph 1210, Engine 1 may demand more power than Engines 2 and 3, which in turn may demand more power than Engine 4.
  • an HV protection function implementing method 1100 for allocating power may determine that total power demanded by the four engines exceeds the HV bus power limit by AP.
  • method 1100 may determine power allocation to each of the components connected to the high-voltage bus based on the difference between the high-voltage bus power limit and the total power demand. Determination of power allocation to each component of the HV bus may involve determining which components may be experiencing power reduction (e.g., power below a normal operating amount, power below a threshold) and how the available power should be allocated to the components of the HV bus.
  • Fig. 13 shows examples of different ways in which HV protection function may determine to allocate power to four engines of a HV bus.
  • Fig. 13 show schematic diagrams illustrating different ways in which power may be allocated to components of the HV bus.
  • Graph A) in Fig. 13 illustrates power demand as indicated by an operation command received by HV protection function, where Engine 1 may require more power than Engines 2 and 3, with Engine 4 demanding the least power.
  • HV protection function may decide that the total power demand based on the operation command exceeds the power limit of the HV bus.
  • the HV protection function may determine an allocation of power that limits power delivery to at least some components connected to the HV bus.
  • allocating power to components of the HV bus may comprise reducing power allocated to multiple (e.g., all) components connected to the high- voltage bus. Reducing power allocated to multiple components connected to the high-voltage bus may include reducing the power proportionally for the multiple components (e.g., all components connected to the high-voltage bus).
  • HV protection function may determine that a solution (e.g., optimized or optimal solution) to respect the HV power limit may include introducing power reduction that is proportional for all engines. Although different engines may demand different amounts of power, HV protection function may determine to reduce power demand of each engine by 25%.
  • method 1100 may include disproportionately reducing power allocated to all components connected to the high-voltage bus.
  • Graph C) of Fig. 13 shows a situation in which the HV protection function, upon determination that power demand by the four engines exceeds the HV bus power limit, may decide to reduce power allocation to each of the four engines in a way that is not proportional to power demand of all engines. For example, power allocated to Engine 1 may be reduced by 25%, while power allocated to Engines 2, 3 and 4 may be reduced by 12%, 15% and 10% respectively.
  • method 1100 may determine non-proportional power reduction based on at least one flight control law (discussed above) or flight control parameter, such as a flight envelope.
  • method 1100 may determine that power demand for at least one first engine at a particular position (e.g., inboard engines) should be prioritized more than power demand for at least one second engine at a different position (e.g., outboard engines), and may reduce power allocation to the at least one second engine more than at least one first engine.
  • a particular position e.g., inboard engines
  • a different position e.g., outboard engines
  • allocating power to components (e.g., each component) connected to the HV bus may include reducing power allocated to at least some components connected to the high-voltage bus, as shown in Graph D) of Fig. 13.
  • HV protection function may determine that power reduction may only be necessary for Engines 1, 2 and 3, which may experience reduction in allocated power by 25%, 15% and 15% respectively.
  • the function may determine that power allocated to Engine 4 may not require altering from power demand associated with the operation command.
  • reducing power allocated to at least some components may be based on a ranking of the components connected to the high-voltage bus.
  • the HV protection function may determine to reduce total power demand to at least some engines by dynamically ranking engines based on real-time performance metrics and/or one or more flight control laws.
  • a performance metric may be associated with (e.g., measure, represent) at least one of efficiency, load, temperature, stability, or reliability.
  • Data from sensors on each engine may be continuously collected and fed into a ranking algorithm that may evaluate and rank the engines.
  • the algorithm may identify one or more engines that may be more efficient, stressed less, and/or prioritized based on a flight control law, and may associate those engines with a higher rank, and the system may then adjust power allocation accordingly. Power may be reduced to the lower-ranked engines while maintaining or increasing power to the higher-ranked ones.
  • engines may be ranked based on their criticality and/or position relative to the aircraft and/or other engines, which may describe the influence of an engine on aircraft’s motion. For example, stability and maneuverability of an aircraft may depend more on output of outboard engines than on the output from engines located closer to the fuselage.
  • the ranking of components connected to the high-voltage bus e.g., engines
  • engines may be ranked based on their criticality with respect to an axis of the aircraft and/or a type of aircraft motion.
  • an outboard engine may be more critical and consequently be ranked higher than other engines for vertical motion, while for execution of pitching motion of the aircraft all engines may have equivalent ranking.
  • HV protection function may reduce power delivered to lower ranked engines in favor of the higher ranked engine.
  • engines may also be ranked based on their current state (e.g., measured by one or more sensors), e.g. temperature, efficiency, vibration level, power consumption, combination thereof, etc. For instance, an engine that is hotter than other engines may also perform less efficiently and consequently, may be ranked lower (e.g., to improve overall power efficiency or reduce stress on the hotter engine, which may be at or near a thermal safety limit) or higher (e.g., to enable the engine to reach than the other engines.
  • their current state e.g., measured by one or more sensors
  • an engine that is hotter than other engines may also perform less efficiently and consequently, may be ranked lower (e.g., to improve overall power efficiency or reduce stress on the hotter engine, which may be at or near a thermal safety limit) or higher (e.g., to enable the engine to reach than the other engines.
  • reducing power allocated to at least some components may comprise at least partial reduction in power delivered to components that are considered non-essential.
  • an air conditioning (AC) unit or an electric anti-ice system e.g., pitot heating system
  • HV protection function may reduce the power allocated to at least one of the AC unit and/or to the electric anti-ice system or may even shut these systems down completely, which may allow for more power to be allocated to other components (e.g., engines).
  • method 1100 may allocate power to multiple components (e.g., each component) connected to the high-voltage bus using the determined power allocation.
  • HV protection function may determine to allocate power as described with reference to Graph B) of Fig. 13, such that all engines experience proportional reduction to power demand. The power may then be distributed to the four engines accordingly.
  • allocating power to at least one engine may comprise determining at least one of torque, power or revolutions per minute (rpm) limit for the at least one engine based on the determined power allocation. Determining power limits based on power, torque or rpm may be beneficial for managing engine performance in an aircraft as these metrics may provide a direct indication of the mechanical output needed for propulsion and lift. This approach may allow for more precise control, ensuring the at least one engine may operate within their optimal efficiency range, preventing excessive mechanical stress and wear. Additionally, implementation of torque, power or rpm limits may improve load sharing, ensuring no single engine may be overburdened, thereby enhancing overall efficiency, safety, and longevity of the at least one engine of the aircraft.
  • rpm revolutions per minute
  • method 1100 may include modifying a command to request a reduced amount of at least one of power, torque or rpm from the at least one engine (e.g., to comply with at least one torque limit of an engine).
  • method 1100 may include determining a difference in power demanded and power allocated for the at least one engine, which may be followed by determination of a torque limit, power limit, rpm limit or any combination thereof for the engine based on the determined difference.
  • the determined torque, power or rpm limit may be utilized to update one or more engine reference models.
  • Updating an engine reference model based on the determined torque, power or rpm limit may involve revising performance curves, control algorithms, and/or safety parameters to adhere to the newly determined limits. This revision may ensure accurate representation of the engine's behavior, safe operation within the new limits, and balanced load distribution across plurality of engines. Additionally, updating engine reference model may include enhancing thermal management, recalibrating sensors and feedback systems, and adjusting maintenance and diagnostic protocols to maintain engine efficiency, reliability, and longevity under the updated operating conditions.
  • determining power allocation may be executed by a high-voltage protection function and the HV protection function may make such determination for multiple received inputs (e.g., every input) indicative of at least one operation command.
  • the execution of HV protection function may be carried out by a flight control computer in the aircraft (e.g. FCC in Fig. 6) while the aircraft is in flight.
  • the HV protection function may be calculated using a closed form algorithm, which may include an algorithm configured to provide an exact solution in a finite number of operations, which may be expressed as a mathematical formula.
  • Fig. 14 shows a schematic diagram illustrating determination of power allocation to components of the HV bus when some components generate power and other components consume power according to embodiments of the present disclosure.
  • HV bus may be connected to six engines with four engines demanding power from the HV bus (Engines 1 - 4) and two engines generating power (Engines 5 and 6).
  • Total power demand may be determined by summing power demand and generation across all engines to calculate total effective power demand (PD, TOTAL) as shown in graph 1420.
  • PD, TOTAL may still exceed the HV bus power limit despite the reduction in total power demand (i.e. the power demand from Engines 1 - 4) by power generation from Engines 4 and 5.
  • the offset between PD, TOTAL and the HV bus power limit AP may still be used to reduce total effective power demand PD, TOTAL as described with reference to Fig. 13.
  • the method for allocating power to components of HV bus may be initiated in response to a fault or failure (e.g., detected by system 1000). Additionally or alternatively, a determined power allocation (e.g., in step 1140) may be based on a fault or failure (e.g., a change in possible power allocations based on component statuses and/or abilities).
  • Fig. 15A illustrates an example of how a fault or a failure may affect power allocation to four engines connected to the HV bus.
  • Graph 1510 shows comparison between total power limit of the HV bus before and after a fault or a failure occurs. In the present example, a fault may result in reduction of available power by the amount equal to AP.
  • Some examples of fault or failures that may result in reduction of the power limit of the HV bus may comprise at least one of a battery failure, power electronics failure, generator malfunction, electrical fault, cooling system failure, connector failure, software glitch, mechanical failure, sensor failure, or control system failure.
  • a consequence of the failure prompt power reduction to electrical components connected to the HV bus, for example, engines.
  • four engines operating at a certain level of power demand prior to occurrence of the fault or failure may experience reduction in power allocation by e.g. 20% each (see graph 1530) in response to loss of total available power by the amount equal to AP.
  • HV protection function may be activated to adjust power allocation to components of HV bus in response to a failure or fault of a component not connected to the HV bus.
  • Fig. 15B illustrates an example of five Engines 1 - 5 connected to a HV bus (HV Bus 1). Initially, the power may be allocated to all five engines as showing in graph 1540.
  • a fault may occur in an engine (Engine X) connected to a different HV bus (HV Bus 2), such that Engine X fails. Failure of Engine X may adversely affect flight of the aircraft in terms of stability and overall safety.
  • HV protection function may respond to the failure of Engine X by changing power allocation to Engines 1 - 5 such that the overall thrust generated by the aircraft satisfies one or more constraints (e.g., resulting in safe and balanced flight). As shown in graph 1550, power allocation may be adjusted such that Engines 1 - 3 may experience power reduction, while Engine 5 may be shut down altogether. Consistent with disclosed embodiments, an engine may be ranked based on its position relative to the aircraft and based on a failure or fault condition.
  • engines closer to it may be ranked higher (e.g., to compensate for its malfunction or disfunction) and engines further from it (e.g., engines on an opposite wing).
  • method 1100 may determine to limit power allocation to a component based on predefined or dynamically determined activation limits.
  • Method 1100 may comprise determining that the total power demand exceeds an activation threshold and activating a high-voltage protection function based on the determination that the total power demand exceeds the activation threshold. Based on the determination that the total power demand exceeds the activation threshold, the high-voltage protection function may be configured to determine the power allocation to components of the HV bus. In some embodiments, the HV protection function may only be deactivated upon detection that the total power demand is below a deactivation threshold.
  • Fig. 16 illustrates an example of activation and deactivation of HV protection function based on total power demand from the HV bus.
  • the total power demand from components connected to the HV bus may vary over time, which is represented by discrete power values at nine different moments in time ti, t2, ts, ... tg.
  • the system e.g., system 1000
  • the system may detect that the total power demand from the received operation input exceeds activation threshold 1610 and in response may activate HV protection function.
  • HV protection function may reduce the total power below threshold 1610.
  • the activation threshold e.g. activation threshold 1610
  • the deactivation threshold e.g.
  • deactivation threshold 1620 At time te Hv protection function may remain active due to power demand still exceeding deactivation threshold 1620. Only at time t-j total power demand drops below deactivation threshold 1620 and the HV protection function may be deactivated. As shown on the example of Fig. 16.
  • deactivation threshold 1620 may be expressed as a proportion of total power limit of the HV bus or activation threshold 1610, or as a mathematical function dependent of total power limit or activation threshold 1610.
  • determining power allocation (e.g., step 1140) may be performed dynamically and the determined power allocation may change over time.
  • the method for allocating power to components of a high-voltage bus of an aircraft as described with reference to Fig. 11 - 16 may also apply to an aircraft comprising multiple HV buses, each bus connected to at least one battery and at least one component.
  • an aircraft may comprise six HV buses, each bus connected to plurality of battery cells and engines.
  • the HV protection function may recognize that there may be six HV buses present in the aircraft and based on the current state of interconnections (e.g. how batteries and engines may be connected to each bus) may group each HV bus with its corresponding batteries and engines. Following identification and grouping of HV buses with their corresponding batteries and engines, the method for allocating power to components of HV bus may be performed for each HV bus and its batteries and/or components as described with reference to Fig. 11 - 16.
  • the method for allocating power to components of a high-voltage bus of an aircraft as described above with reference to Fig. 11 - 16 may be implemented into a system, e.g. high-voltage power system (HVPS) 610 in Fig. 6.
  • the system may comprise a high-voltage bus and/or at least one processor configured to carry out the method for allocating power to components of a high-voltage bus of an aircraft.
  • the method described above with reference to any or all of Fig. 11 - 16 may be implemented in an aircraft, for example aircraft 100 of Fig. 1.
  • the aircraft may comprise fuselage 102, at least one wing (e.g.
  • wing 104) mounted to fuselage 102, at least one stabilizer mounter to the rear of the fuselage (e.g. stabilizer 106) and a plurality of propellers mounted to the at least one wings (e.g. propellers 114), where at least one of the propellers may be tiltable.
  • a system e.g. HVPS 610 in Fig. 6 which may be configured to carry out a method for allocating power to components of a high-voltage bus of an aircraft as described above with reference to any or all of Fig. 11 - 16. In some embodiments, the method for allocating power to components of a high-voltage bus of an aircraft as described above with reference to any or all of Fig.
  • 11 - 16 may be stored in a non-transitory computer readable medium in form of instructions that may be executed by at least one processor of a system (e.g., HVPS 610 in Fig. 6, FCS 612 in Fig. 6, an FCC in Fig. 8) to carry out the method.
  • a system e.g., HVPS 610 in Fig. 6, FCS 612 in Fig. 6, an FCC in Fig. 8
  • a method for allocating power to components of a high-voltage bus of an aircraft comprising: receiving an input indicative of an operation command; calculating, based on the received input, a total power demand associated with a plurality of components connected to the high-voltage bus, each component having a respective component power demand; comparing the total power demand to a high-voltage bus power limit assigned to the high-voltage bus to determine a difference between the high-voltage bus power limit and the total power demand; determining power allocation to each of the components connected to the high-voltage bus based on the difference between the high-voltage bus power limit and the total power demand; and allocating power to each component connected to the high-voltage bus using the determined power allocation.
  • determining the high-voltage bus power limit assigned to the high-voltage bus is based on at least one of measured state of charge of the battery cells, measured state of power of the battery cells or measured state of health of the battery cells.
  • allocating power comprises reducing power allocated to all components connected to the high-voltage bus.
  • allocating power comprises reducing power allocated to at least some components connected to the high-voltage bus.
  • reducing power allocated to at least some components comprises: reducing power based on a ranking of the components connected to the high-voltage bus.
  • allocating the power includes modifying a command to request a reduced amount of at least one of power, torque or revolutions per minute from the at least one engine.
  • a vertical take-off and landing (VTOL) aircraft comprising: a fuselage; at least one wing mounted to the fuselage; at least one stabilizer mounter to the rear of the fuselage; a plurality of propellers mounted to the at least one wing, wherein at least one of the propellers is tiltable; a high-voltage bus electrically connected to a plurality of battery cells; and a flight control system comprising at least one processing unit configured to carry out the method of any of clauses 1 - 23.
  • VTOL vertical take-off and landing
  • a system for allocating power to components of a high-voltage bus of an aircraft comprising: a high-voltage bus; at least once processor configured to: receive an input indicative of an operation command; calculate, based on the received input, a total power demand associated with a plurality of components connected to the high-voltage bus, each component having a respective component power demand; compare the total power demand to a high-voltage bus power limit assigned to the high-voltage bus to determine a difference between the high-voltage bus power limit and the total power demand; determine power allocation to each of the components connected to the high- voltage bus based on the difference between the high-voltage bus power limit and the total power demand; and allocate power to each component connected to the high-voltage bus using the determined power allocation.
  • the at least one processor is further configured to: determine a difference in power demand and power allocated for the at least one engine; determine at least one of power limit, torque limit or revolutions per minute limit for the engine based on the determined difference; and update an engine reference model based on the determined at least one of power limit, torque limit, or revolutions per minute limit.
  • the at least one processor is further configured to: determine that the total power demand exceeds an activation threshold; activate a power allocation function based on the determination that the total power demand exceeds the activation threshold, wherein the power allocation function is configured to determine the power allocation; and deactivate the power allocation upon detection that the total power demand is below a deactivation threshold.

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Abstract

L'invention concerne des systèmes et des procédés de commande de vol d'aéronefs entraînés par des systèmes de propulsion électrique et d'autres types de véhicules, comprenant un procédé d'attribution de puissance à des composants d'un bus à haute tension d'un aéronef. Le procédé consiste à recevoir une entrée indiquant une instruction d'opération ; calculer, sur la base de l'entrée reçue, une demande de puissance totale associée à une pluralité de composants connectés au bus à haute tension, chaque composant ayant une demande de puissance de composant respective ; comparer la demande de puissance totale à une limite de puissance de bus à haute tension attribuée au bus à haute tension pour déterminer une différence entre la limite de puissance de bus à haute tension et la demande de puissance totale ; déterminer une attribution de puissance à chacun des composants connectés au bus à haute tension sur la base de la différence entre la limite de puissance de bus à haute tension et la demande de puissance totale ; et attribuer une puissance à chaque composant connecté au bus à haute tension à l'aide de l'attribution de puissance déterminée.
PCT/US2024/037420 2023-07-10 2024-07-10 Systèmes et procédés pour protections dynamiques de canaux de puissance Pending WO2025015072A1 (fr)

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Citations (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5936318A (en) * 1996-05-03 1999-08-10 Daimlerchrysler Aerospace Airbus Gmbh Power distribution arrangement especially in an aircraft
US20200307390A1 (en) * 2019-03-25 2020-10-01 Beta Air Llc Systems and methods for maintaining attitude control under degraded energy source conditions using multiple propulsors

Patent Citations (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5936318A (en) * 1996-05-03 1999-08-10 Daimlerchrysler Aerospace Airbus Gmbh Power distribution arrangement especially in an aircraft
US20200307390A1 (en) * 2019-03-25 2020-10-01 Beta Air Llc Systems and methods for maintaining attitude control under degraded energy source conditions using multiple propulsors

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