System and method for controlling an electric vertical take-off and landing vehicle

Abstract

A method of controlling an electric aircraft, the electric aircraft including a plurality of actuators, the actuators including a plurality of electric propulsion units, the method comprising: receiving force and torque commands for the electric aircraft; determining control commands for the plurality of actuators based on desired force and torque commands by solving an optimization problem, the optimization problem including a noise minimization term for minimizing noise generated by the electric propulsion units; and controlling the plurality of actuators according to the determined control commands to satisfy the force and torque commands for the electric aircraft.

Description

[0001] CROSS-REFERENCE TO RELATED APPLICATIONS

[0002] This application claims priority to and the benefit of U.S. Application No. 17 / 157,580, filed on January 25, 2021, which is hereby incorporated by reference in its entirety. Technical Field

[0003] The present invention relates generally to aircraft control, and in particular to the control of electric vertical take-off and landing aircraft. Background Art

[0004] A vertical take-off and landing (VTOL) aircraft is an aircraft capable of vertical takeoff and landing and hovering, providing the ability to transport travelers directly to their destinations. A helicopter is a VTOL aircraft that generates lift entirely through its rotors. Some VTOL aircraft have wings and a propulsion system that enables the wings to provide the lift required during forward flight. Some winged VTOL aircraft use separate propulsion systems for vertical thrust during takeoff and landing and for forward thrust during cruising. Other winged VTOL aircraft use a tiltable propulsion system that tilts between vertical and forward thrust positions. Electric VTOL aircraft use electric propulsion units to provide thrust for both vertical and forward flight. Many electric VTOL aircraft include movable electric propulsion units, where the thrust vector of the propulsion units can be changed, for example, from an upward direction for vertical lift to a forward direction for forward flight. Because there are more actuator degrees of freedom than motion degrees of freedom, many electric VTOL aircraft are over-actuated. Control distribution refers to the issue of distributing control forces among the multiple actuators in an over-actuated system. Electric VTOL aircraft typically include more propulsion units and other actuators than traditional aircraft, and the propulsion units and other actuators strongly affect multiple control axes. Therefore, electric VTOL aircraft may face a greater control distribution problem than traditional aircraft. Summary of the Invention

[0005] According to some embodiments, systems and methods for control allocation in an electric VTOL aircraft include considering rotor acoustics when controlling the aircraft's actuators. According to various embodiments, control allocation includes solving an optimization objective function that includes satisfying force and torque commands as a primary objective and modulating rotor acoustics as a secondary objective. In some embodiments, rotor acoustics are modulated by minimizing edgewise flight, varying rotor speeds in a rotor group, and / or minimizing propeller tip speed.

[0006] According to some embodiments, systems and methods for control allocation include considering battery pack charge when controlling an aircraft's actuators. In some embodiments, an electric VTOL aircraft includes multiple battery packs that are electrically isolated from one another and power one or more electric propulsion units. Controlling the allocation may include solving an optimization objective function that includes satisfying force and torque commands as a primary objective and balancing energy usage across the battery packs as a secondary objective. In some embodiments, electric propulsion units powered by battery packs with greater charge are preferentially utilized.

[0007] According to various embodiments, a method of controlling an electric aircraft, the electric aircraft including a plurality of actuators including a plurality of electric propulsion units, the method comprising: receiving force and torque commands for the electric aircraft; determining control commands for the plurality of actuators based on desired force and torque commands by solving an optimization problem, the optimization problem including a noise minimization term that minimizes noise generated by the electric propulsion units; and controlling the plurality of actuators according to the determined control commands to satisfy the force and torque commands for the electric aircraft.

[0008] In any of these embodiments, controlling the multiple actuators according to the determined control command may include: causing at least a first electric propulsion unit among the multiple electric propulsion units to operate at a speed different from that of at least a second electric propulsion unit among the multiple electric propulsion units to disperse the frequencies of the multiple electric propulsion units over a wider frequency band.

[0009] In any of these embodiments, electric propulsion units closer to the fuselage of the aircraft may be operated at a lower speed than electric propulsion units further from the fuselage to reduce noise at the fuselage.

[0010] In any of these embodiments, the electric propulsion units may operate at different speeds during straight-forward flight.

[0011] In any of these embodiments, at least a portion of the plurality of electric propulsion units may be tiltable, and controlling the plurality of actuators according to the determined control commands may include tilting the electric propulsion units and adjusting at least one of the attitude of the aircraft to minimize edgewise flight time.

[0012] In any of these embodiments, controlling the plurality of actuators according to the determined control commands may include setting a pitch of blades of at least one electric propulsion unit to minimize a speed of the at least one electric propulsion unit.

[0013] In any of these embodiments, the electric aircraft may be a vertical take-off and landing aircraft.

[0014] In any of these embodiments, the electric aircraft can be manned.

[0015] In any of these embodiments, the electric aircraft may include a plurality of electric propulsion units on either side of the aircraft's fuselage.

[0016] According to various embodiments, a system for controlling an electric aircraft includes a plurality of actuators, the plurality of actuators including a plurality of electric propulsion units, the system including one or more processors, a memory, and one or more programs, the one or more programs stored in the memory for execution by the one or more processors to: receive desired force and torque commands for the electric aircraft; determine control commands for the plurality of actuators based on the desired force and torque commands by solving an optimization problem, the optimization problem including a noise minimization term that minimizes noise generated by the electric propulsion units; and control the plurality of actuators according to the determined control commands to meet the desired force and torque commands for the electric aircraft.

[0017] According to various embodiments, a method for controlling an electric aircraft, wherein the electric aircraft includes a plurality of actuators, wherein the plurality of actuators include a plurality of electric propulsion units and a plurality of battery packs for powering the plurality of electric propulsion units, the method comprising: receiving desired force and torque commands for the electric aircraft; monitoring an energy state of the plurality of battery packs, wherein at least a first battery pack of the plurality of battery packs is electrically isolated from at least a second battery pack of the plurality of battery packs; determining control commands for the plurality of actuators based on the desired force and torque commands by solving an optimization problem, the optimization problem comprising energy balance terms for balancing energy draw of the electric propulsion units according to the monitored energy states of the plurality of battery packs; and controlling the plurality of actuators according to the determined control commands to meet the desired force and torque commands for the electric aircraft.

[0018] In any of these embodiments, the first battery pack has a lower remaining energy than the second battery pack, and a first electric propulsion unit powered by the first battery pack operates at a lower power than a second electric propulsion unit powered by the second battery pack.

[0019] In any of these embodiments, the first battery pack and the second battery pack have the same energy capacity.

[0020] In any of these embodiments, the first electric propulsion unit and the second electric propulsion unit have the same power rating.

[0021] In any of these embodiments, the energy balance item includes a set of preferred operating states for the plurality of electric propulsion units, and the preferred operating state for an electric propulsion unit powered by a battery pack with lower residual energy is lower than the preferred operating state for an electric propulsion unit powered by a battery pack with higher residual energy.

[0022] In any of these embodiments, the energy balance term includes a set of penalties for deviating from the preferred operating state, and the penalty associated with an electric propulsion unit connected to a lower energy battery pack is higher than the penalty associated with an electric propulsion unit connected to a higher energy battery pack.

[0023] In any of these embodiments, the optimization problem includes a noise minimization term that minimizes noise generated by the electric propulsion unit.

[0024] In any of these embodiments, the electric aircraft is a vertical take-off and landing aircraft.

[0025] In any of these embodiments, the electric aircraft is manned.

[0026] In any of these embodiments, the electric aircraft includes a plurality of electric propulsion units on either side of a fuselage of the aircraft.

[0027] According to various embodiments, a system for controlling an electric aircraft, the electric aircraft including a plurality of actuators, the plurality of actuators including a plurality of electric propulsion units, the system including one or more processors, a memory and one or more programs, the one or more programs being stored in the memory for execution by the one or more processors: receiving desired force and torque commands for the electric aircraft; monitoring an energy state of the plurality of battery packs, wherein at least a first battery pack of the plurality of battery packs is electrically isolated from at least a second battery pack of the plurality of battery packs; determining control commands for the plurality of actuators based on the desired force and torque commands by solving an optimization problem, the optimization problem including energy balance terms balancing the energy draw of the electric propulsion units according to the monitored energy states of the plurality of battery packs; and controlling the plurality of actuators according to the determined control commands to meet the desired force and torque commands for the electric aircraft. BRIEF DESCRIPTION OF THE DRAWINGS

[0028] The present invention will now be described, by way of example only, with reference to the accompanying drawings, in which:

[0029] Figure 1A shows a VTOL aircraft in a forward flight configuration according to various embodiments;

[0030] Figure 1Bshows a VTOL aircraft in a take-off and landing configuration according to various embodiments;

[0031] Figure 2A and Figure 2B shows a power distribution architecture for powering an electric propulsion unit of an aircraft according to various embodiments;

[0032] Figure 3 is a block diagram of a method for control distribution for an electric aircraft according to some embodiments;

[0033] Figure 4 is a functional block diagram of a control system for controlling actuators of an electric aircraft according to various embodiments;

[0034] Figure 5 is a functional block diagram of a control distribution system including multi-step optimization with frequency partitioning according to various embodiments; and

[0035] Figure 6 An example of a computing system 700 is shown in accordance with various embodiments. DETAILED DESCRIPTION

[0036] According to various embodiments, systems and methods for electric VTOL aircraft control distribution include considering rotor acoustics and / or battery pack energy balance when distributing control forces among actuators of the aircraft. According to various embodiments, control distribution includes solving an optimization objective function that includes achieving force and torque commands as a primary objective and modulating rotor acoustics and / or balancing battery pack energy as secondary objectives.

[0037] According to various embodiments, an electric VTOL aircraft includes multiple electric propulsion units on each side of the aircraft fuselage. The electric propulsion units include rotor-driven propellers and can reduce the acoustic noise generated by the aircraft by operating the rotors at different speeds to spread the combined frequencies over a larger frequency band, which can reduce the amplitude of any individual frequency, resulting in lower perceived noise. In some embodiments, the rotor acoustics can be modulated by minimizing the propeller tip speed while still providing the required thrust. This can be achieved by determining the rotor speed and propeller pitch that achieves the lowest tip speed for the required thrust. In some embodiments, at least a portion of the electric propulsion units are tiltable, and the rotor acoustics can be modulated by utilizing thrust steering and flight trajectories to minimize edgewise flight time. In some embodiments, the rotor acoustics can be modulated by balancing the thrust distribution to minimize excessive demand on any one rotor.

[0038] According to various embodiments, an electric VTOL aircraft includes a plurality of electrically isolated battery packs to provide power to different portions of a plurality of electric propulsion units. In some embodiments, control allocation includes battery pack energy balancing, which can be achieved by minimizing the use of electric propulsion units connected to battery packs with lower charges than other battery packs. According to various embodiments, minimizing the use of electric propulsion units connected to battery packs with lower charges can be achieved by reducing the preferred state of the electric propulsion units in a control allocation optimization objective function and / or increasing the penalty for deviations from the preferred state.

[0039] In the following description of the present invention and embodiments, reference is made to the accompanying drawings, in which are shown by way of illustration specific embodiments that can be practiced. It should be understood that other embodiments and examples can be practiced and changes can be made without departing from the scope of the present invention.

[0040] In addition, it should be understood that the singular forms "a", "an" and "the" used in the following description are also intended to include the plural forms, unless the context clearly indicates otherwise. It should also be understood that the term "and / or" as used herein refers to and encompasses any and all possible combinations of one or more of the associated listed items. It should also be understood that the terms "comprises", "comprising", "includes" and / or "comprising" when used herein specify the presence of the stated features, numbers, steps, operations, elements, parts and / or units, but do not preclude the presence or addition of one or more other features, numbers, steps, operations, elements, parts, units and / or combinations thereof.

[0041] As used herein, the term "proprotor" refers to a variable pitch propeller that can provide thrust for vertical lift and forward propulsion by varying the pitch of the propeller.

[0042] As used herein, the term "battery pack" means any combination of electrically connected batteries (ie, battery cells), and may include multiple batteries arranged in series, parallel, or a combination of series and parallel.

[0043] Figure 1A and Figure 1B The VTOL aircraft 100 is shown in a cruise configuration and a vertical take-off and landing configuration, respectively, according to various embodiments. Exemplary embodiments of a VTOL aircraft according to various embodiments are discussed in U.S. patent application Ser. No. 16 / 878,380, filed May 19, 2020, entitled “Vertical Take-Off and Landing Aircraft,” which is incorporated herein by reference in its entirety.

[0044] Aircraft 100 includes a fuselage 102, wings 104 mounted to fuselage 102, and one or more rear stabilizers 106 mounted to the rear of fuselage 102. Aircraft 100 includes a plurality of rotors 112 and a plurality of propulsion rotors 114 (collectively referred to herein as EPUs). The EPUs (112, 114) typically include electric motors that drive a plurality of blades and motor controllers for controlling / powering the motors. In some embodiments, the pitch of the blades of one or more of the EPUs can be controlled in flight. In some embodiments, the EPUs can include multiple local motors that can drive the blades independently and collectively and can be controlled by multiple separate motor controllers.

[0045] Rotors 112 are mounted to wing 104 and are configured to provide lift for vertical takeoff and landing. Propeller rotors 114 are mounted to wing 104 and are tiltable between a lift configuration and a propulsion configuration. In the lift configuration, propulsion rotors 114 provide a portion of the lift required for vertical takeoff and landing and hovering, as shown in FIG. Figure 1B As shown, in the propulsion configuration, propulsion rotors 114 provide forward thrust to aircraft 100 for horizontal flight, as shown in FIG. Figure 1A As used herein, a lift configuration of a propulsive rotor refers to any propulsive rotor orientation in which the thrust of the propulsive rotor primarily provides lift to the aircraft, and a thrust configuration of a propulsive rotor refers to any propulsive rotor orientation in which the thrust of the propulsive rotor primarily provides forward thrust to the aircraft.

[0046] According to various embodiments, rotors 112 are configured to provide lift only, with all thrust provided by the propulsion rotors. Thus, rotors 112 can be in a fixed position. During takeoff and landing, propulsion rotors 114 are tilted to a lift configuration, in which their thrust is directed downward to provide additional lift. In some embodiments, rotors 112 are tiltable for thrust diversion.

[0047] For forward flight, the propulsion rotors 114 tilt from their lift configuration to their propulsion configuration. In other words, the tilt of the propulsion rotors 114 changes from a range of tilted positions in which the thrust of the propulsion rotors is directed upward to provide lift during vertical takeoff and landing and during hovering to a range of tilted positions in which the thrust of the propulsion rotors is directed forward to provide forward thrust to the aircraft 100. The propulsion rotors tilt about an axis 118 that is perpendicular to the forward direction of the aircraft 100. When the aircraft 100 is in full forward flight, lift can be provided entirely by the wings 104 and the rotors 112 can be closed. The blades 120 of the rotors 112 can be locked in a low drag position for aircraft cruising. In some embodiments, each rotor 112 has two blades 120 that are locked in a minimum drag position for cruising in which one blade is directly in front of the other blade, as in FIG. Figure 1A In some embodiments, rotor 112 has more than two blades. In some embodiments, propulsion rotor 114 includes more blades 116 than rotor 112. For example, Figure 1A and Figure 1B As shown, rotors 112 may each include two blades, while propulsion rotors 114 may each include five blades. According to various embodiments, propulsion rotors 114 may have from two to five blades.

[0048] According to various embodiments, the aircraft includes only one wing 104 on each side of fuselage 102 (or a single wing extending across the entire aircraft), and at least a portion of rotors 112 is located behind wings 104, and at least a portion of propulsion rotors 114 is located in front of wings 104. In some embodiments, all rotors 112 are located behind wings 104, and all propulsion rotors are located in front of wings 104. According to some embodiments, all rotors 112 and propulsion rotors 114 are mounted to the wings, i.e., no rotors or propulsion rotors are mounted to the fuselage. According to various embodiments, all rotors 112 are located behind wings 104, and all propulsion rotors 114 are located in front of wings 104. According to some embodiments, all rotors 112 and propulsion rotors 114 are located inboard of wingtips 109.

[0049] According to various embodiments, rotor 112 and propulsion rotor 114 are mounted to wing 104 via boom 122. Boom 122 can be mounted below wing 104, on top of the wing, and / or can be integrated into the airfoil. According to various embodiments, one rotor 112 and one propulsion rotor 114 are mounted to each boom 122. Rotor 112 can be mounted at the rear end of boom 122, while propulsion rotor 114 can be mounted at the front end of boom 122. In some embodiments, rotor 112 is mounted in a fixed position on boom 122. In some embodiments, propulsion rotor 114 is mounted to the front end of boom 122 via hinge 124. Propulsion rotor 114 can be mounted to boom 122 such that propulsion rotor 114 is aligned with the main body of boom 122 when in its propulsion configuration, forming a continuous extension of the front end of boom 122 that minimizes drag in forward flight.

[0050] According to various embodiments, the aircraft 100 may include multiple wings on each side of the aircraft 100, only one wing on each side of the aircraft 100, or a single wing extending across the aircraft 100. According to some embodiments, at least one wing 104 is a high wing mounted to the upper side of the fuselage 102. According to some embodiments, the wing includes control surfaces 150, such as flaps and / or ailerons, which are positioned via one or more control surface actuators (not shown). According to some embodiments, the wing may have a curved wingtip 109 for reducing drag during forward flight. According to some embodiments, the rear stabilizer 106 includes control surfaces 152, such as one or more rudders, one or more elevators, and / or one or more combined rudder-elevator, which are positioned via one or more control surface actuators (not shown). The wing(s) may have any suitable design. In some embodiments, the wing has a tapered leading edge 123, such as at Figure 1A In some embodiments, the wing has a tapered trailing edge.

[0051] Figure 2A A power distribution architecture for powering the EPUs (112, 114) of the aircraft 100 is shown in accordance with various embodiments. Figures 1A to 2A 12 EPUs ( Figure 2A 1 to 12), but it can have any suitable number of EPUs, including 4, 6, 8, 10, 14, 18, 20 or more. The EPUs are powered by a plurality of battery packs 200. Figure 2A In the embodiment shown, there are six battery packs 200, numbered 1 to 6. Each battery pack 200 only powers a portion of the EPU. In the embodiment shown, each battery pack 200 powers two EPUs. Figure 2B The following are listed in Figure 2A The illustrated embodiment groups battery packs and EPUs. Battery pack 1 powers EPU 1 and EPU 12, battery pack 2 powers EPU 2 and EPU 11, and so on. Each battery pack 200 is connected to its corresponding EPU via a dedicated power distribution bus (e.g., buses 202 and 204). Therefore, the power distribution bus 202 of battery pack 1 is not electrically connected to the power distribution bus 204 of battery pack 2.

[0052] Because the battery packs 200 are electrically isolated from each other, an electrical failure in one pack or its power distribution circuit does not affect the operation of other EPUs and packs. Only the EPU powered by the failed pack or distribution circuit is affected. Consequently, there is no single point of failure in the aircraft's power supply. Furthermore, because the battery packs and power distribution circuits are isolated from each other, no diodes are required to prevent current from flowing from one pack to another. This can significantly save weight and improve efficiency compared to systems with parallel battery packs.

[0053] According to various embodiments, the particular EPUs powered by a given battery pack may be selected to reduce the destabilizing effects caused by the loss of power to the EPUs in the event of a battery pack failure. According to various embodiments, EPUs arranged on opposite sides of one or more axes of symmetry of a set of EPUs may be powered by the same battery pack to reduce roll, pitch, or yaw moments that may be caused by the loss of power to the EPUs driven by the battery packs. For example, EPUs located in the same relative position on either side of the longitudinal axis 280 of the aircraft may be driven by a first battery pack such that if one of the battery packs fails, minimal roll moment will occur because the thrust provided by the remaining EPUs will remain uniform about the longitudinal axis. Similarly, in some embodiments, a group of EPUs are arranged at least partially forward of the leading edges of a pair of wings, and a group of EPUs are arranged at least partially rearward of the trailing edges of a pair of wings, and the EPUs on opposite sides of the wings and opposite sides of the longitudinal axis 280 may be powered by the same battery pack such that minimal roll and pitch moments (such as those in FIG. 2 ) will occur in the event of a battery pack failure. Figure 2A shown).

[0054] According to various embodiments, each battery pack 200 provides power to at least a portion of at least one propulsion rotor 114 and at least a portion of at least one rotor 112. Figure 2A In the embodiment of FIG, the rotors and propulsion rotors in opposite positions are driven by the same battery pack 200. Therefore, the outermost propulsion rotor 114 ( Figure 2A The EPU1 in the figure is composed of the same battery pack ( Figure 2A Similarly, the other pair of outermost EPUs ( Figure 2A The EPUs 6 and 7 in FIG are powered by the same battery pack (Battery Pack 6). Grouping need not be limited to EPUs in precise relative positions. For example, EPU 1 can be grouped with EPU 11 but not with EPU 12.

[0055] The number of EPUs powered by a given battery pack can be greater than two. For example, in some embodiments, the number of EPUs per battery pack can be three, four, five, six, or any other suitable fraction of the total number of EPUs. According to various embodiments, there can be a different number of EPUs within each group. For example, one group can have two EPUs (two EPUs powered by one battery pack), while another group can have four EPUs (four EPUs powered by a different battery pack). The number of battery packs can be as few as two. In various embodiments, the number of battery packs is at least three, at least four, at least five, at least six, at least seven, at least eight, or more.

[0056] Figure 3 is a block diagram of a method 300 for control distribution for an electric aircraft according to some embodiments. The method 300 may be used for an electric aircraft (e.g. Figure 1A Method 300 determines commands for a plurality of actuators of the aircraft. The plurality of actuators may include one or more actuators associated with each of a plurality of electric propulsion units of the aircraft, and may include one or more control surface actuators of the aircraft. As discussed further below, method 300 may include a control distribution that utilizes over-actuation of the aircraft to optimize rotor acoustic noise and / or battery pack energy balance. According to some embodiments, method 300 utilizes additional degrees of freedom gained due to the number and configuration of the electric propulsion units to reduce acoustic noise while meeting the forces and moments desired for the aircraft. According to some embodiments, the aircraft includes a plurality of battery packs that independently power the electric propulsion units, and method 300 utilizes the additional degrees of freedom to balance the battery pack energy.

[0057] At step 302, force and torque commands for an electric aircraft are received, which may include receiving force and torque commands from an upstream computing module of the aircraft (e.g., a flight control system), which may generate force and torque commands based on pilot input. The force and torque commands may be desired forces and torques to be applied to the aircraft through the action of one or more of the various actuators of the aircraft.

[0058] As used herein, an "actuator" is any subsystem of an aircraft that provides a degree of freedom for controlling the aircraft. For example, the rotor of each EPU is an actuator, and its degree of freedom is the rotational speed of the blades. One or more EPUs may include other actuators, such as a rotor pitch system to provide thrust steering and / or a propeller blade pitch system for adjusting the pitch of the propeller blades. Thus, for example, Figure 1A and Figure 1BThe aircraft 100 may include up to three actuators (rotor speed, blade pitch, and rotor pitch) on each propulsion rotor 114 and up to two actuators (rotor speed, blade pitch) on each rotor 112, which for the EPU configuration shown can provide up to 30 actuators for the EPU (24 actuators for embodiments where rotor blade pitch is not adjustable). Other actuators may be included on the wings of the aircraft (e.g., Figure 1A control surfaces 150) and / or tail (e.g. Figure 1A and Figure 1B In some embodiments, there are 10 control surface actuators, and the total number of actuators available for control distribution according to method 300 is 34.

[0059] At step 304, control commands for at least a portion of the aircraft's actuators are determined by solving an optimization problem that includes minimizing a cost function that seeks to satisfy the force and torque commands from step 302 and achieve one or more secondary objectives that may include reducing sound generated by the aircraft's EPU and / or balancing energy usage of multiple battery packs powering the electric propulsion unit. The control commands include control commands for the actuators of the aircraft's EPU, which may include, for example, rotor speed, propeller blade pitch, and / or rotor pitch (e.g., for a propulsion rotor).

[0060] According to some embodiments, one or more secondary objectives include modulating the noise generated by the electric propulsion unit of the aircraft, e.g. Figure 3 , as shown in optional block 306 in the . According to various embodiments, noise reduction can be achieved by reducing the tip speed of the propulsion rotors of one or more EPUs. According to some embodiments, the tip speed of the EPU's propellers can be reduced by adjusting the propeller pitch while providing the required thrust. In some embodiments, noise reduction can be achieved by distributing the load across the EPUs to reduce the peak load at any one EPU.

[0061] In some embodiments, noise reduction can be achieved by operating the EPUs at different speeds so that the sound frequencies produced by the EPUs are spread over a wider frequency band that is perceived as less noise by the human ear. In some embodiments, EPUs closer to the fuselage operate at lower speeds than EPUs further away from the fuselage to reduce the amount of noise generated closer to the fuselage and its passengers. In some embodiments, the spread of frequencies is mirrored along the centerline of the aircraft so that EPUs in the same position on either side of the aircraft operate at the same speed to maintain thrust balance, such as during direct flight. For example, Figure 1AThe two innermost propulsive rotors 114 on either side of the aircraft 100 of the example in can be operated at the slowest speed of all the propulsive rotors, while the two outermost propulsive rotors 114 on either side of the aircraft 100 can be operated at the highest speed of all the propulsive rotors while maintaining a balance of thrust about the yaw axis.

[0062] In some embodiments, at least a portion of the EPUs are tiltable, and rotor noise acoustics can be modulated by utilizing thrust steering to minimize edgewise time. The cost function in step 304 can assign a relatively high cost to operating one or more of the tiltable EPUs in a full vertical thrust position, such that, when force and torque commands permit, minimization of the cost function will tend to move the tiltable EPUs away from their full vertical thrust position faster than strictly required to meet the force and torque commands. According to some embodiments, thrust steering to reduce acoustic noise can be offset by attitude adjustments actuated via control surfaces.

[0063] According to some embodiments, one or more secondary objectives include balancing the energy draw of the electric propulsion units based on the energy states of multiple battery packs of the aircraft, as shown in optional block 308 of method 300. In some embodiments, the aircraft includes multiple battery packs that independently power different EPUs and can be electrically isolated from each other. Via the control allocation of step 304, the energy states of the battery packs can be monitored and the EPUs can be controlled so that an EPU powered by a lower energy battery pack can be used less than an EPU powered by a higher energy battery pack. For example, with respect to Figure 2A and Figure 2B In the illustrated embodiment, at some point during flight, battery pack 1 may have a relatively lower charge than battery pack 2, and to balance the energy of the battery packs, according to various embodiments, one or more EPUs powered by battery pack 1 may operate at a relatively lower power state, and one or more EPUs connected to battery pack 2 may operate at a relatively higher power state to make up for at least a portion of the lost thrust from the one or more EPUs connected to battery pack 1.

[0064] In some embodiments, a relatively low charge can mean a lower overall charge, such as when the battery packs have the same energy capacity and one battery pack has a lower remaining charge than the other battery pack. In some embodiments, a relatively low charge can be relative to the capacity of the corresponding battery pack. For example, Figure 2A and Figure 2B The capacity of battery pack 2 may be lower than the capacity of battery pack 1, but its relative charge may be higher than the relative charge of battery pack 1 (eg, 90% for battery pack 2 and 80% for battery pack 1).

[0065] In some embodiments, one or more EPUs connected to a lower energy battery pack operate at a lower power than one or more EPUs connected to a higher energy battery pack. Figure 2A and Figure 2B , at least one EPU connected to battery pack 1 and at least one EPU connected to battery pack 2 may have the same rated power, and the EPU connected to battery pack 1 may operate at a lower power than the EPU connected to battery pack 2. In some embodiments, one or more EPUs connected to the lower-energy battery pack operate at a lower power relative to the nominal power for a given aircraft operating state, which may or may not be a lower power than one or more EPUs connected to the higher-energy battery pack, which operate at a higher power relative to the nominal power for a given aircraft operating state. For example, an EPU connected to the lower-energy battery pack may have a higher rated power than an EPU connected to the higher-energy battery pack, and the EPU connected to the lower-energy battery pack may operate at a relatively lower power that is still higher than the relatively higher power of the EPU connected to the higher-energy battery pack.

[0066] At step 310, at least a portion of the vehicle's actuators are operated in accordance with the control commands determined in step 304 to meet the desired force and torque commands for the electric vehicle. For example, various actuators associated with a plurality of electric propulsion units are operated in accordance with the determined control commands. Depending on the secondary objectives sought in the optimization of step 304 (which may include modulating EPU acoustic noise, battery pack energy balance, or both), step 310 may achieve not only the desired forces and torques on the vehicle, but also relatively low acoustic noise and / or a certain degree of energy balance in the vehicle's battery pack.

[0067] Figure 4is a functional block diagram of a control system 400 for controlling actuators of an electric VTOL aircraft, such as aircraft 100, in accordance with various embodiments. The system 400 includes a control allocation module 402 that generates actuator commands 404 based on various inputs 406. As discussed further below, the control allocation module 402 determines the actuator commands 404 by minimizing an objective function that includes one or more primary objectives (e.g., satisfying commanded aircraft forces and moments) and one or more secondary objectives (which may include minimizing acoustic noise and / or optimizing battery pack usage). The control system 400 may be implemented by a microprocessor-based controller executing software code stored in a storage medium to perform the functions described herein. The control system 400 may also be implemented in hardware or a combination of hardware and software. The control system 400 may be implemented as part of a flight control system for an aircraft. It should be understood that for ease of description, many of the conventional functions of the control system are not described herein. Figure 4 Shown in.

[0068] Inputs 406 to the control distribution module 402 may include one or more of the following: force and torque commands 408, actuator states 410, envelope protection limits 412, scheduling parameters 414, aerodynamic parameters 416, battery state 418, and optimizer parameters 420. The force and torque commands 408 include up to six force and torque commands, which may include x, y, and z force commands and x, y, and z torque commands. As is known in the art, the force and torque commands may be derived from operator commands (or autopilot commands, or commands from an autonomous controller of an unmanned aerial vehicle) and aircraft state (e.g., velocity, acceleration, altitude, attitude). The force and torque commands 408 are generated by an upstream controller (not shown) and provided to the control distribution module 402.

[0069] The actuator status 410 includes actuator hardware limitations, such as travel limits, speed limits, response time limits, etc., and may include an actuator health indicator that may indicate actuator performance degradation, which may limit the ability of a given actuator to meet actuator commands. The actuator status 410 may be used to determine limits (e.g., minimum / maximum values) for various actuator commands. According to various embodiments, the battery status 418 is the remaining energy in the aircraft's battery pack, which may be monitored when controlling distribution, including balancing battery pack energy states.

[0070] Envelope protection limits 412 may include commanded limits that prevent operation outside of the flight envelope, which defines the operating limits of the aircraft, including limits based on velocity and acceleration as known in the art.

[0071] Scheduling parameters 414 are velocity-related parameters used to define the allocation problem. Aerodynamic parameters 416 are parameters obtained from aerodynamic and acoustic modeling and can be based on the actuator Jacobian matrix and the actuator state. Aerodynamic parameters 416 can be a function of scheduling parameters 414. Optimizer parameters 420 are parameters used to define the optimization problem, as discussed further below. Optimizer parameters 420 can include axis weights that define the relative priority of force and torque axes. Optimizer parameters 420 can also include individual actuator weights that define the relative importance of different actuators in the control allocation problem. In some embodiments, optimizer parameters 420 are a function of scheduling parameters 414.

[0072] The control allocation module 402 may include a limit calculation module 430, a parameter interpolation module 432, and an optimization module 434. The limit calculation module 430 calculates limits for each actuator command based on the actuator state 410 and the envelope protection limits 412. In normal operation, the minimum command limit for a given actuator comprises the maximum of the following: the minimum hardware-based limit and the minimum flight envelope limit; and the maximum command limit for a given actuator comprises the minimum of the following: the maximum hardware-based limit and the maximum flight envelope limit. In the event of an actuator failure, the command limit for the failed actuator corresponds to the failure mode (e.g., the position of an unresponsive control surface actuator or zero RPM for a failed rotor).

[0073] The parameter interpolation module 432 can be configured to determine parameters that vary with the scheduling parameters, which are based on the speed of the aircraft, as described above. The value of a given parameter (e.g., an aerodynamic parameter or an optimizer parameter) can be determined from a lookup table of parameters based on the scheduling parameter associated with the current speed of the aircraft, as follows:

[0074] x 输出 =F(x 表格 ,v)

[0075] in:

[0076]

[0077] In some embodiments, the parameter interpolation module 432 can determine one or more parameters associated with one or more secondary objectives (e.g., battery pack energy balance). For example, the parameters for battery pack energy balance can be determined from a lookup table of parameters based on the battery state 418 and the scheduling parameters, such as as follows:

[0078] x 输出 =F(x 表格 ,v,E 电池 )

[0079] in,

[0080]

[0081]

[0082] Optimization module 434 executes a nonlinear optimization algorithm that minimizes the sum of a primary objective 440 and a secondary objective 442. Primary objective 440 seeks to satisfy force and torque commands 408 and may prioritize commands by axis in the event of saturation. In the event that the solution space for forces and torques includes multiple actuator command combinations, secondary objective 442 seeks to satisfy other operational objectives. Secondary objectives 442 may include terms for modulating rotor acoustics and / or balancing energy usage, as discussed further below. Other operational objectives that may be included in secondary objectives 442 may include prioritizing specific actuators and minimizing deviations from one or more preferred actuator states.

[0083] The optimization module 434 minimizes an objective function that includes a primary objective 440 and a secondary objective 442. The following is an example of an objective function:

[0084]

[0085] According to various embodiments, the optimization module 434 seeks a set of actuator commands u that minimizes the objective function of Expression 1, which conforms to:

[0086] min≤u≤max

[0087] In Expression 1, the first function Corresponds to the primary objective 440. W comprises weights used to prioritize specific force and moment axes and is determined by the interpolation module 432 based on the optimizer parameters 420. B is the actuator Jacobian matrix, which is determined by the interpolation module 432 based on the aerodynamic parameters. FM req are the force and torque commands.

[0088] The second function in expression 1 corresponds to the secondary objective 442 and comprises the null space of the Jacobian matrix B. ε comprises weights for the relative priorities of the actuators and is determined by the interpolation module 432 based on the optimizer parameters 420. u0 comprises the actuator preferred operating state and is determined by the interpolation module 432 based on the aerodynamic parameters 416.

[0089] The secondary objective function in Expression 1 is the l2 norm to minimize the deviation from the preferred operating state of the actuator. Other minimization functions can be used, such as using the l1 norm to minimize the control force or using l ∞ norm to minimize the maximum order.

[0090] According to various embodiments, energy balancing of multiple electric propulsion units powered by electrically isolated battery packs is included in the secondary objective by adjusting at least one of the weight ε and the preferred operating state u0 of the actuators based on the battery state 418. For example, the weight ε of the actuators associated with the lower energy battery pack can have a higher value (for a higher penalty for deviating from the preferred state) than the weight ε of the actuators associated with the higher energy battery pack. Additionally or alternatively, the preferred operating state u0 of the actuators associated with the lower energy battery pack can have a lower value than the preferred operating state u0 of the actuators associated with the lower energy battery pack.

[0091] According to various embodiments, acoustic modulation of the plurality of electric propulsion units of an aircraft is included in a secondary objective by applying preferred operating states u0 that minimize the sound received in the cabin. For example, at any given aircraft speed, the preferred operating speeds of the rotors and / or propulsion rotors are dispersed over a frequency range according to the method outlined in

[00039] in order to reduce the noise perceived by passengers.

[0092] The optimization module 434 seeks the set of actuator commands u that minimizes the objective function of Expression 1, subject to the actuator command constraints u determined by the constraint calculation module 430. 最小 with u 最大 Requirements between.

[0093] According to various embodiments, the objective function may be solved by formulating the objective function as a quadratic problem and solving the quadratic problem using a quadratic programming solver. Examples of suitable quadratic programming solvers include interior point, active set, conjugate gradient, and augmented Lagrangian solvers.

[0094] Figure 5 is a functional block diagram of a system 500 for control distribution including a multi-step optimization with frequency partitioning according to various embodiments. The control distribution of the system 500 includes a first control distribution step performed based on low-frequency force and torque commands and a second control distribution step performed based on high-frequency force and torque commands. The system 500 can ensure that high-frequency commands are not generated for slow actuators that cannot respond to high-frequency commands. Similar to Figure 4 Compared to system 400 , system 500 may provide a control allocation that seeks to achieve one or more secondary objectives (eg, reduced noise and / or battery pack energy balance) with lower computational cost relative to a single-step optimization.

[0095] In system 500, multiple inputs 502 are provided to a force and torque command filter 508. Multiple inputs 502 include force and torque commands 504 and may include optimizer parameters 506. Filter 508 filters low-frequency force and torque commands 510 from unfiltered force and torque commands 504. Filter 508 may filter the low-frequency force and torque commands based on a filter cutoff frequency and gain, which may be predetermined parameters based on scheduling parameters. The dynamic characteristics and / or power consumption of the actuators may be used to determine the filter behavior. For example, the actuator response time may be used to determine an appropriate filter cutoff frequency so that relatively low-frequency actuator commands are generated for relatively slow actuators (i.e., actuators with relatively long response times). According to various embodiments, the group of slow actuators may include an actuator for controlling the pitch of an electric propulsion unit and an actuator for controlling the pitch of a blade, while the group of fast actuators may include an electric propulsion unit motor and one or more control surface actuators for controlling blade speed. According to various embodiments, the actuator states, envelope protection limits, and / or scheduling parameters may be fed directly into the slow actuation allocation module 512 to determine limits for the slow actuator commands 514. According to various embodiments, the actuator states, envelope protection limits, and / or scheduling parameters may be fed directly into the fast actuation allocation module 518 to determine limits for the fast actuator commands 524. Other inputs, such as interpolated aerodynamic parameters and / or battery state, may also be fed directly into the modules 512 and 518 to form the objective functions solved in the respective modules.

[0096] The low frequency force and torque commands 510 are provided to a slow actuation distribution module 512 which can be used in a manner similar to Figure 4 The optimization module 434 is configured in a manner such that a set of actuator commands is determined by minimizing an objective function, as discussed above with respect to the control distribution module 402. A set of slow actuator commands 514 is output from the slow actuation distribution module 512 to control the slow actuators.

[0097] The slow actuation distribution module 512 may also output a set of fast actuator commands 516, which may be provided to a fast actuation distribution module 518. The fast actuation distribution module 518 may also be provided with a set of high-frequency force and torque commands 520. The high-frequency force and torque commands 520 may be provided to the fast actuation distribution module 518 by subtracting from the force and torque commands 504 the forces and torques 522 that the slow actuation distribution module 512 has determined and implemented from the slow actuator commands 514 and the fast actuator commands 516 determined by the slow actuation distribution module 512.

[0098] The fast actuation allocation module 518 determines a set of actuator commands by minimizing an objective function, as discussed above with respect to the control allocation module 402, and in a manner similar to the slow actuation allocation module 512. The fast actuation allocation module 518 outputs only the fast actuator commands 524 for controlling the fast actuators. Thus, the slow actuators are controlled based on the slow actuator commands 514 from the slow actuation allocation module 512, and the fast actuators are controlled based on the fast actuator commands 524 from the fast actuation allocation module 518.

[0099] According to various embodiments, when the frequency of the force and torque commands 504 is higher than the slow actuators alone can handle, the slow actuation allocation module 512 will output slow actuator commands 514 that can approach, but not exceed, the hardware limitations of the slow actuators. The unfulfilled portion of the force and torque commands 504 can then be satisfied by the fast actuators via the fast actuation allocation module 518. While the multi-step optimization with frequency partitioning of the system 500 is computationally more intensive than the single-step optimization control allocation, it allows for the simultaneous utilization of both slow and fast actuators.

[0100] Figure 6 An example of a computing system 600 is shown, which may be used in the system 400 of FIG. 1 and / or Figure 5 6. One or more components of system 500, such as control dispensing module 402 of system 400, slow actuation dispensing module 512 of system 500, and / or fast actuation dispensing module 518 of system 500. System 600 can be any suitable type of processor-based system. System 600 can include, for example, one or more of the following: an input device 620, an output device 630, one or more processors 610, a memory 640, and a communication device 660.

[0101] Input device 620 may be any suitable device that allows user input (e.g., user input from a pilot), such as one or more buttons, joysticks and / or switches, one or more touch screens, etc. Output device 630 may be or include any suitable device that provides output, such as a display, a touch screen, a haptic device, a virtual / augmented reality display, or a speaker.

[0102] Memory 640 may be any suitable device that provides storage, such as electrical, magnetic, or optical storage including RAM, cache, a hard drive, a removable storage disk, or other non-transitory computer-readable medium. Communication device 660 may include any suitable device or combination of devices capable of sending and receiving signals from one or more other computing systems or modules. The components of computing system 600 may be connected in any suitable manner, such as via a physical bus connection or wirelessly.

[0103] The processor 610 may be any suitable processor or combination of processors, including any one or any combination of a central processing unit (CPU), a graphics processing unit (GPU), a field programmable gate array (FPGA), and an application specific integrated circuit (ASIC). The software 650 that may be stored in the memory 640 and executed by one or more processors 610 may include, for example, programs that implement the functionality or portion of the functionality of the present invention (e.g., as implemented in the above-mentioned devices). For example, the software 650 may include one or more programs executed by one or more processors 610 to perform one or more steps of the methods described herein, such as Figure 3 One or more steps of method 300.

[0104] The software 650 may also be stored and / or transmitted on any non-transitory computer-readable storage medium for use by or in conjunction with an instruction execution system, apparatus, or device, such as those described above, from which the software 650 may retrieve instructions associated with the software and execute the instructions. In the context of the present invention, a computer-readable storage medium may be any medium, such as memory 640, that may contain or store a program for use by or in conjunction with an instruction execution system, apparatus, or device.

[0105] The software 650 may also be transmitted over any transmission medium for use by or in conjunction with an instruction execution system, device, or apparatus, such as those described above, from which the software 650 may retrieve instructions associated with the software and execute the instructions. In the context of the present invention, a transmission medium may be any medium capable of communicating, propagating, or transmitting a program for use by or in conjunction with an instruction execution system, device, or apparatus. Transmission computer-readable media may include, but are not limited to, electrical, magnetic, optical, electromagnetic, or infrared wired or infrared wireless propagation media. The system 600 may implement any suitable operating system and may be written in any suitable programming language.

[0106] For purposes of illustration, the foregoing description has been described with reference to specific embodiments. However, the foregoing illustrative discussion is not intended to be exhaustive or to limit the invention to the precise forms disclosed. In light of the above teachings, many modifications and variations are possible. The embodiments were chosen and described in order to best explain the principles of the technology and its practical application. Thus, others skilled in the art will be able to best utilize the technology and various embodiments with various modifications to suit the particular application contemplated.

[0107] Although the present invention and examples have been fully described with reference to the accompanying drawings, it should be noted that various changes and modifications will become apparent to those skilled in the art. Such changes and modifications should be understood to be included within the scope of the present disclosure and examples as defined by the claims. Finally, the entire disclosures of the patents and publications referred to in this application are hereby incorporated by reference.

Claims

1. A method of controlling an electric aircraft, the electric aircraft comprising a plurality of actuators, the plurality of actuators comprising a plurality of electric propulsion units, the method comprising: receiving desired force and torque commands for the electric aircraft; determining control commands for the plurality of actuators based on desired force and torque commands by solving an optimization objective function, the optimization objective function including a noise minimization term for minimizing noise generated by the electric propulsion unit; as well as controlling the plurality of actuators according to the determined control commands to satisfy force and torque commands for the electric aircraft, Wherein, at least a portion of the plurality of electric propulsion units are tiltable, and controlling the plurality of actuators according to the determined control command comprises at least one of tilting the electric propulsion units and adjusting the attitude of the aircraft to minimize edgewise flight time.

2. The method according to claim 1, wherein Controlling the plurality of actuators according to the determined control commands includes operating at least a first electric propulsion unit of the plurality of electric propulsion units at a speed different from that of at least a second electric propulsion unit of the plurality of electric propulsion units to disperse frequencies of the plurality of electric propulsion units over a wider frequency band.

3. The method according to claim 2, wherein: Electric propulsion units closer to a fuselage of the electric aircraft are operated at a lower speed than electric propulsion units further away from the fuselage to reduce noise at the fuselage.

4. The method according to claim 2, wherein: The electric propulsion units operate at different speeds during straight-forward forward flight.

5. The method according to any one of claims 1 to 4, wherein Controlling the plurality of actuators according to the determined control commands includes setting a pitch of blades of at least one electric propulsion unit to minimize a speed of the at least one electric propulsion unit.

6. The method according to any one of claims 1 to 4, wherein The electric aircraft is a vertical take-off and landing aircraft.

7. The method according to any one of claims 1 to 4, wherein The electric aircraft is manned.

8. The method according to any one of claims 1 to 4, wherein The electric aircraft includes a plurality of electric propulsion units on either side of a fuselage of the electric aircraft.

9. A system for controlling an electric aircraft, the electric aircraft comprising a plurality of actuators, the plurality of actuators comprising a plurality of electric propulsion units, the system comprising: one or more processors, Memory, and one or more programs stored in the memory for execution by the one or more processors: receiving desired force and torque commands for the electric aircraft; determining control commands for the plurality of actuators based on the desired force and torque commands by solving an optimization objective function, the optimization objective function including a noise minimization term for minimizing noise generated by the electric propulsion unit; as well as controlling the plurality of actuators according to the determined control commands to meet desired force and torque commands for the electric aircraft, Wherein, at least a portion of the plurality of electric propulsion units are tiltable, and controlling the plurality of actuators according to the determined control command comprises at least one of tilting the electric propulsion units and adjusting the attitude of the aircraft to minimize edgewise flight time.

10. The system according to claim 9, wherein: Controlling the plurality of actuators according to the determined control commands includes operating at least a first electric propulsion unit of the plurality of electric propulsion units at a speed different from that of at least a second electric propulsion unit of the plurality of electric propulsion units to disperse frequencies of the plurality of electric propulsion units over a wider frequency band.

11. The system according to claim 10, wherein: Electric propulsion units closer to a fuselage of the electric aircraft are operated at a lower speed than electric propulsion units further away from the fuselage to reduce noise at the fuselage.

12. The system according to claim 10, wherein: The electric propulsion units operate at different speeds during straight-forward forward flight.

13. The system according to claim 10, wherein: Controlling the plurality of actuators according to the determined control commands includes setting a pitch of blades of at least one electric propulsion unit to minimize a speed of the at least one electric propulsion unit.

14. The system according to any one of claims 10 to 13, wherein: The electric aircraft is a vertical take-off and landing aircraft.

15. The system according to any one of claims 10 to 13, wherein: The electric aircraft is manned.

16. The system according to any one of claims 10 to 13, wherein: The electric aircraft includes a plurality of electric propulsion units on either side of the aircraft's fuselage.

17. An electric aircraft, comprising: a plurality of actuators, the plurality of actuators comprising a plurality of electric propulsion units; and A system according to any one of claims 10 to 13 for controlling the electric aircraft via the plurality of actuators.

18. The electric aircraft according to claim 17, wherein: The electric aircraft includes a plurality of electric propulsion units on either side of a fuselage of the electric aircraft.

19. The electric aircraft according to claim 17, wherein: The electric aircraft is a vertical take-off and landing aircraft.

20. The electric aircraft according to claim 17, wherein: The electric aircraft is manned.

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