Electrically powered vertical take-off and landing aircraft

Abstract

To provide a VTOL aircraft comprising a plurality of motor assemblies, each configured to generate thrust by moving air through the motor assembly along each axis of thrust of the motor assembly, and a wing.SOLUTION: The VTOL aircraft is such that: 1) orientations of axes of thrust are each fixed, during operation of the aircraft, at a constant pitch angle oblique to a pitch orientation of the wing; 2) a plurality of motor assemblies are operable together to both fully support the aircraft in a hovering mode and propel the aircraft forward in a forward flight mode; 3) the wing does not intersect with any right cylinder centered on any motor assembly, having a central longitudinal axis aligned with the axis of thrust of the motor assembly, and having a radius equal to a radius of a propeller of the motor assembly.SELECTED DRAWING: Figure 1A

Description

[Technical Field]

[0001] This application claims the benefit of priority under 35 U.S.C. § 119(e) of U.S. Provisional Patent Application No. 62 / 786,564, filed December 31, 2018, the contents of which are incorporated herein by reference in their entirety.

[0002] The present invention relates to the field of airborne flying vehicles, and more particularly to electric aircraft with vertical take-off and landing as well as geostationary flight capabilities. [Background technology]

[0003] Multicopter aircraft designs use multiple power-driven propellers to provide vertical lift and / or horizontal thrust. Multicopters are used and proposed for a range of applications, including the transport of cargo and / or passengers.

[0004] Practical multicopters are made possible by at least one of the following technologies: computerized control and / or sensing electronics; relatively light and powerful electric motors; improved battery capacity-to-weight ratios; and improved power generation using relatively light generators.

[0005] An advantage of multicopters is that they require little or no horizontal speed to be generated, especially before takeoff. Lift generated by fixed wings efficiently converts forward momentum into vertical thrust that can keep the aircraft aloft. Some built and proposed aircraft designs combine the capability for vertical takeoff and / or landing using propeller-generated thrust with lift generated by fixed wings in forward flight. Summary of the Invention [Means for solving the problem]

[0006] According to some aspects of the present disclosure, there is provided an aircraft including a plurality of motor assemblies, each configured to generate thrust by moving air past the motor assembly along a respective thrust axis of the motor assembly, the directions of the thrust axes being fixed at constant respective pitch angles oblique to the pitch direction of the wing during operation of the aircraft, the plurality of motor assemblies together operable to both fully support the aircraft in a hover mode and to propel the aircraft forward in a forward flight mode, and each motor assembly is positioned within a radius of the motor assembly in a position such that the wing does not intersect with the airflow before or after entering or exiting the motor assembly in a direction along the respective thrust axis.

[0007] According to some embodiments of the present disclosure, at speeds above 55 km / h generated by thrust from the multiple motor assemblies, the wing is configured to provide at least 25% of the lift required to support the weight of the aircraft in forward flight.

[0008] According to some embodiments of the present disclosure, at speeds generated by thrust from the multiple motor assemblies, the wings provide at least 50% of the lift required to support the aircraft against gravitational acceleration in forward flight.

[0009] According to some embodiments of the present disclosure, the thrust axis is oriented approximately 55° to 80° away from the pitch direction of the wing.

[0010] According to some embodiments of the present disclosure, the thrust axis is approximately 65° to 70° from the pitch direction of the wing. °Facing away.

[0011] According to some embodiments of the present disclosure, the wing is disposed along the roll axis of the aircraft between at least two of the plurality of motor assemblies.

[0012] According to some embodiments of the present disclosure, at least one of the plurality of motor assemblies is positioned forward of the center of gravity of the aircraft and below a wing, and at least one of the plurality of motor assemblies is positioned aft of the center of gravity of the aircraft and above a wing.

[0013] According to some embodiments of the present disclosure, the plurality of motor assemblies includes at least one motor assembly mounted to the aircraft on either side of the aircraft's center of gravity along the aircraft's roll axis.

[0014] According to some embodiments of the present disclosure, the multiple motor assemblies are operable to fully support the air vehicle during ground-stationary hovering of the air vehicle.

[0015] According to some embodiments of the present disclosure, the multiple motor assemblies are operable to provide horizontal thrust to the aircraft while fully supporting the weight of the aircraft.

[0016] According to some embodiments of the present disclosure, the multiple motor assemblies are operable to provide horizontal thrust to the aircraft while fully supporting the aircraft against downward acceleration due to gravity and while the aircraft is pitched forward with the wings at their most efficient angle of attack.

[0017] According to some embodiments of the present disclosure, each of the multiple motor assemblies is mounted at the same oblique pitch angle relative to the pitch direction of the blade.

[0018] According to some embodiments of the present disclosure, at least two of the plurality of motor assemblies are mounted within the span of the wing.

[0019] According to some embodiments of the present disclosure, the wings are configured to provide less than 100% of the lift required to support the weight of the aircraft during forward flight.

[0020] According to some embodiments of the present disclosure, the aircraft is configured to rotate its pitch angle from a hovering orientation using thrust from multiple motor assemblies to a forward flight orientation where the wings are at their most efficient angle of attack, while the motor assemblies fully support the aircraft against downward acceleration due to gravity.

[0021] According to some embodiments of the present disclosure, the aircraft is configured to rotate its pitch angle by tilting the nose of the aircraft upward from a landing orientation, in which the aircraft is stationary on level ground, to a hovering orientation using thrust from multiple motor assemblies.

[0022] According to some embodiments of the present disclosure, the aircraft is configured to rotate its pitch angle from a hovering orientation using thrust from the multiple motor assemblies to a forward flight orientation in which the wings are at their most efficient angle of attack by tilting the nose of the aircraft downward.

[0023] According to some embodiments of the present disclosure, the aircraft is sized to carry at least one occupant.

[0024] According to some embodiments of the present disclosure, the motor assembly includes a power source and a rotor that is rotated by the power source. and a rotor including a propeller.

[0025] According to some embodiments of the present disclosure, the motor assembly is mounted to the aircraft through the fuselage.

[0026] According to some embodiments of the present disclosure, the oblique pitch angle between the thrust axis and the pitch direction of the wing is less than 75°.

[0027] According to some embodiments of the present disclosure, the power source is an electric motor coaxially aligned with the propeller.

[0028] According to some embodiments of the present disclosure, the oblique pitch angle between the thrust axis and the pitch direction of the wing is greater than 45°.

[0029] According to some embodiments of the present disclosure, when the aircraft is stationary on the ground, the contact surface between the landing gear and the ground is oblique to both the pitch direction of the wing and the thrust axis of the multiple motor assemblies.

[0030] According to an aspect of some embodiments of the present disclosure, there is provided a method of launching an aircraft into forward flight from the ground, the method including: starting with the aircraft in a fully grounded position, applying thrust from a motor assembly of the aircraft at an oblique pitch angle relative to the ground; redirecting thrust from the motor assembly of the aircraft to a pitch angle normal to the ground while maintaining contact with the ground; and redirecting thrust from the motor assembly obliquely relative to the ground surface to accelerate the aircraft horizontally, the motor assembly being mounted to a frame of the aircraft, and each reorientation including reorienting the frame of the aircraft.

[0031] According to some embodiments of the disclosure, the aircraft includes fixed wings that are fixed in orientation for at least 50% of the mass of the aircraft, and includes rotating the wings with rotation of the aircraft from a fully grounded pitch angle position oblique to its horizontal flight direction, to a pitch angle further oblique to its horizontal flight direction while the aircraft is airborne, and then to the horizontal flight direction of the aircraft during forward motion of the aircraft.

[0032] In accordance with an aspect of some embodiments of the present disclosure, there is provided a self-tilting rotor mounting assembly for an aircraft that includes: a pivot bar connected to the aircraft and configured to pivot relative to the aircraft about a pivot axis parallel to a pitch axis of the aircraft; and a plurality of rotors mounted on opposite sides of the pivot axis of the pivot bar, the rotors operable to rotate the bar about the pivot axis.

[0033] According to some embodiments of the present disclosure, the pivot angle range of the pivot bar relative to the aircraft is mechanically limited between more horizontal angles that direct thrust from the rotor in a more vertical direction and more vertical angles that direct thrust from the rotor in a more horizontal direction.

[0034] According to some embodiments of the present disclosure, the pivot axis is located aft of the aircraft's center of gravity and the aircraft's center of aerodynamic lift, and is configured so that the weight of the aircraft holds the pivot bar at a more horizontal angle during hovering flight, and aerodynamic lift reduces the weight of the aircraft during forward motion of the aircraft, allowing the pivot bar to rotate in a pitch downward direction without tilting the aircraft downward.

[0035] According to some embodiments of the present disclosure, the pivot is configured to have a resistance that slows the rotation when the rotor operates to rotate the bar about the pivot axis.

[0036] In accordance with one aspect of some embodiments of the present disclosure, a variable blade pitch rotor is provided, comprising: first and second electric motors, each comprising a stator and a rotor, each electric motor being arranged coaxially relative to one another; and a plurality of propeller blades attached to each rotor of the first and second electric motors, respectively, wherein the pitch of the propeller blades is varied according to the relative angular positions of the rotors of the first and second electric motors.

[0037] According to some embodiments of the present disclosure, a change in the relative angular position of the rotors induces one of the following: rotation of an individual gear to which each propeller blade is attached; and movement of an individual lever to which each propeller blade is attached.

[0038] According to some embodiments of the present disclosure, the pivot bar rotates in a pitch downward direction relative to the aircraft to switch between a more horizontal angle and a more vertical angle.

[0039] According to some embodiments of the present disclosure, the propeller blade pitch is configured to adjust the blade pitch from minimum pitch to maximum pitch in less than one second.

[0040] According to one aspect of some embodiments of the present disclosure, there is provided a protected propeller for a winged aircraft, including a protection body surrounding the propeller in the plane of rotation of the propeller, the protection body being shaped like an oblique cylinder having a top side and a bottom side, with leading and trailing side walls of the protection body oriented approximately parallel to the wing surface of the aircraft from the top to the bottom, and both lateral walls of the protection body oriented obliquely to the wing surface of the aircraft from the top to the bottom.

[0041] According to some embodiments of the present disclosure, the leading and trailing sides of the protector are shaped with an airfoil cross section, each having a relatively more blunted leading edge and a relatively more tapered trailing edge.

[0042] According to an aspect of some embodiments of the present disclosure, there is provided an electric motor including a rotor and a stator, the stator including a plurality of individual winding coils spaced apart about the periphery of the stator, each of the individual winding coils being wound by an individual wire that is not wound around any of the other coils and occupying a single location about the periphery of the stator, and all of the individual winding coils being configured to simultaneously receive electrical power acting on the rotor.

[0043] According to some embodiments of the present disclosure, a single location of individually wound coils is not internally separated into multiple portions by other ones of the individually wound coils.

[0044] According to some embodiments of the present disclosure, the electric motor includes an individual controller for each individual winding coil configured to select at least one of a magnitude and a polarity of a current to supply to the coil according to a relative position of the rotor.

[0045] According to an aspect of some embodiments of the present disclosure, a battery-powered aircraft is provided that includes a fuselage surrounding a center of gravity of the aircraft, a plurality of battery units, and a plurality of pairs of electric drive rotors, the rotors of each pair being mounted on diagonally opposite sides of the center of gravity and separated along a periphery of the center of gravity by at least one other rotor of the plurality of pairs, and each battery unit powers both rotors of one of the plurality of pairs of electric drive rotors.

[0046] According to some embodiments of the present disclosure, each of the pair of electric drive rotors is powered by one or more of the plurality of battery units.

[0047] According to an aspect of some embodiments of the present disclosure, there is provided a multi-rotor air vehicle that includes a fuselage, a plurality of rotors mounted to the fuselage, and a plurality of independently operating flight controller units, each including a respective inertial measurement unit (IMU), wherein each rotor is independently controlled by a respective flight control unit based on measurements from the respective IMU of the flight control unit.

[0048] According to some embodiments of the present disclosure, each of the multiple flight controller units receives measurements from an individual IMU of each of the other control units.

[0049] According to some embodiments of the present disclosure, the flight control units are interconnected such that each has access to the same estimate of the overall aircraft flight conditions.

[0050] According to some embodiments of the present disclosure, each individual winding coil is configured to act as a rotor position sensor, and an individual controller supplies current to the coil based on sensor measurements from the individual winding coil.

[0051] According to some embodiments of the present disclosure, each flight controller unit issues commands aimed at producing the same overall aircraft flight state based on the same estimate of the overall aircraft flight state.

[0052] According to some embodiments of the present disclosure, each flight controller unit is configured to issue commands to the same estimate of flight conditions for the entire aircraft, modified by changes in its own IMU measurements, including adjustments for flexing of the multi-rotor aircraft frame.

[0053] In accordance with an aspect of some embodiments of the present disclosure, there is provided a multi-rotor air vehicle including a fuselage, a plurality of rotors mounted to the fuselage, each rotor including a plurality of counter-rotating coaxially mounted propellers, and a flight controller configured to exercise yaw authority over the air vehicle by adjusting the rotational speed of only one of the respective propellers of the plurality of rotors.

[0054] According to some embodiments of the present disclosure, each of the multiple rotors is tilted to direct a portion of the rotor thrust in a yaw direction.

[0055] According to some embodiments of the present disclosure, the flight controller is configured to simultaneously adjust the yaw thrust and yaw torque such that each adjustment results in inducing yaw in the same direction.

[0056] According to an aspect of some embodiments of the present disclosure, there is provided an aircraft including a plurality of motor assemblies, each configured to generate thrust by moving air past the motor assembly along a respective thrust axis of the motor assembly, and a wing, wherein the plurality of motor assemblies together are operable to fully support the aircraft in a hover mode and to propel the aircraft forward in a forward flight mode, the plurality of motor assemblies being mounted within a span of the wing and positioned such that when each motor assembly operates at a pitch angle oblique to the pitch direction of the wing, each motor assembly does not cross, within a radius of the motor assembly, the airflow before or after entering or exiting the motor assembly along the thrust axis.

[0057] According to some embodiments of the present disclosure, the multiple motor assemblies are located forward and below the wing. and a motor assembly mounted aft and above the wing.

[0058] In accordance with an aspect of some embodiments of the present disclosure, there is provided a multi-rotor air vehicle operable in both a hover mode and a forward flight mode, the multi-rotor air vehicle including a fuselage, an aerodynamic surface attached to the fuselage including at least horizontal lift-generating surfaces and vertically projecting yaw stabilizing surfaces, a plurality of rotors attached to the fuselage, and a flight controller, the flight controller configured to exercise continuous yaw authority to stabilize the air vehicle by regulating the rotational speed of the rotors in the hover mode, and also configured to cease regulating the rotational speed of the rotors to stabilize the air vehicle yaw in the forward flight mode.

[0059] According to one aspect of some embodiments of the present disclosure, there is provided an aircraft including a plurality of motor assemblies, each configured to generate thrust by moving air rearwardly along a respective thrust axis of the motor assembly, and a wing, wherein the aircraft has a mass of at least 50 kg, and during operation of the aircraft, the directions of the thrust axes are fixed at respective constant pitch angles oblique to the pitch direction of the wing, and the plurality of motor assemblies are operable together to both fully support the aircraft in a hover mode and to propel the aircraft forward in a forward flight mode.

[0060] According to one aspect of some embodiments of the present disclosure, there is provided a method of setting a thrust vector of a motor assembly of an aircraft, the method including providing a propeller having peak efficiency for a given thrust while moving vertically through the air at a selected airspeed, planning a cruise speed flight period for the aircraft above the selected airspeed, and setting the thrust vector of the motor assembly at an angle selected to recover at least a portion of the efficiency lost by moving above the selected speed.

[0061] Unless otherwise defined, all technical and scientific terms used herein have the same meaning as commonly understood by one of ordinary skill in the art to which this disclosure belongs. Although methods and materials similar or equivalent to those described herein can be used in the practice or testing of embodiments of the present disclosure, at least one exemplary method and material is described below. In case of conflict, the present specification, including definitions, will control. Furthermore, the materials, methods, and examples are illustrative only and are not intended to be necessarily limiting.

[0062] As will be appreciated by those skilled in the art, aspects of the present disclosure may be embodied as a system, method, or computer program product. Accordingly, aspects of the present disclosure may take the form of an entirely hardware form, an entirely software form (including firmware, resident software, microcode, etc.), or an embodiment combining software and hardware aspects, all of which may be collectively referred to herein as a "circuit," "module," or "system" (e.g., a method may be implemented using "computer circuitry"). Furthermore, some embodiments of the present disclosure may take the form of a computer program product embodied in one or more computer-readable medium(s) having computer-readable program code embodied therein. Implementation of the methods and / or systems of some embodiments of the present disclosure may involve performing and / or completing selected tasks manually, automatically, or a combination thereof. Furthermore, depending on the actual instrumentation and implementation of at least some embodiments of the methods and systems of the present disclosure, some selected tasks may be implemented by hardware, software, firmware, and / or a combination thereof, for example, using an operating system.

[0063] For example, hardware for performing selected tasks according to some embodiments of the present disclosure may be implemented as a chip or circuit. As software, selected tasks according to some embodiments of the present disclosure may be implemented as a plurality of software instructions executed by a computer using any suitable operating system. In some embodiments of the present disclosure, one or more tasks performed in the method and / or system are performed by a data processor (also referred to herein as a "digital processor," in reference to a data processor that operates with groups of digital bits), such as a computing platform that executes a plurality of instructions. Optionally, the data processor includes at least one of volatile memory and non-volatile storage for storing instructions and / or data, e.g., at least one of a magnetic hard disk and a removable medium for storing instructions and / or data. Optionally, a network connection is also provided. Optionally, a display and / or at least one of an input device, such as a keyboard or a mouse, are also provided. Any of these implementations are more generally referred to herein as an instance of computer circuitry.

[0064] Any combination of one or more computer-readable media can be used for some embodiments of the present disclosure. The computer-readable medium can be a computer-readable signal medium or a computer-readable storage medium. The computer-readable storage medium can be, for example, but not limited to, an electronic, magnetic, optical, electromagnetic, infrared, or semiconductor system, apparatus, or device, or any suitable combination thereof. More specific examples (non-exhaustive list) of computer-readable storage media can include an electrical connection having one or more wires, a portable computer diskette, a hard disk, a random access memory (RAM), a read-only memory (ROM), an erasable programmable read-only memory (EPROM or flash memory), an optical fiber, a portable compact disc read-only memory (CD-ROM), an optical storage device, a magnetic storage device, or any suitable combination thereof. In the context of this specification, a computer-readable storage medium can be any tangible medium that can hold or store a program for use by or in connection with a system, apparatus, or device that executes instructions. A computer-readable storage medium is capable of holding or storing information for use by such programs, such as data recorded by the computer-readable storage medium in a structure accessible by the computer program, such as one or more tables, lists, arrays, data trees, and / or other data structures. Herein, a computer-readable storage medium that records data in a form that can be read as groups of digital bits is also referred to as digital memory. It should be understood that in some embodiments, a computer-readable storage medium is optionally also used as a computer-writable storage medium, provided that the computer-readable storage medium is not inherently read-only and / or is not in a read-only state.

[0065] As used herein, a data processor is said to be "configured" to perform data processing operations so long as the data processor is coupled to a computer-readable memory so as to receive instructions and / or data therefrom, process them, and / or store the results of the processing in the same or another computer-readable storage memory. The processing (optionally on the data) to be performed is specified by the instructions. The processing operations may additionally or alternatively be referred to by one or more other terms, such as at least one of comparing, evaluating, determining, calculating, identifying, associating, storing, analyzing, selecting, and transforming. For example, in some embodiments, a digital processor receives instructions and data from a digital memory, processes the data in accordance with the instructions, and / or stores the results of the processing in the digital memory. In some embodiments, "providing" the results of the processing includes at least one of transmitting, storing, and presenting the results of the processing. Presenting optionally includes displaying the results. This includes displaying, sounding, printing, or otherwise rendering the results in a form accessible to human perception.

[0066] A computer-readable signal medium may include a propagated data signal having computer-readable program code embodied therein, for example in baseband or as part of a carrier wave. Such a propagated signal may take any of a variety of forms, including but not limited to, electromagnetic, optical, or any suitable combination thereof. A computer-readable signal medium is not a computer-readable storage medium and may be any computer-readable medium that can communicate, propagate, or transport a program for use by or in connection with an instruction execution system, apparatus, or device.

[0067] The program code embodied on the computer-readable medium and / or data used thereby may be transmitted using any suitable medium, including but not limited to wireless, wired, fiber optic cable, RF, etc., or any suitable combination thereof.

[0068] Computer program code for carrying out operations of some embodiments of the present disclosure may be written in any combination of one or more programming languages, including object-oriented programming languages ​​such as Java, Smalltalk, C++, etc., and conventional procedural programming languages ​​such as the "C" programming language or similar programming languages. The program code may run entirely on the user's computer, partially on the user's computer as a stand-alone software package, partially on the user's computer and partially on a remote computer, or entirely on a remote computer or server. In the latter case, the remote computer may be connected to the user's computer via any type of network, including a local area network (LAN) or a wide area network (WAN), or may be connected to an external computer (e.g., via the Internet using an Internet Service Provider).

[0069] Some embodiments of the present disclosure may be described below with reference to flowchart illustrations and / or block diagrams of methods, apparatus (systems), and computer program products according to embodiments of the present disclosure. It will be understood that each block of the flowchart illustrations and / or block diagrams, and combinations of blocks in the flowchart illustrations and / or block diagrams, can be implemented by computer program instructions. These computer program instructions are provided to a processor of a general-purpose computer, a special-purpose computer, or other programmable data processing apparatus forming a machine, such that the instructions, executed by the processor of the computer or other programmable data processing apparatus, generate means for performing the functions / acts identified in one or more blocks of the flowchart illustrations and / or block diagrams.

[0070] These computer program instructions can also be stored on a computer-readable medium that can instruct a computer, other programmable data processing device, and / or other device to function in a particular manner, and the instructions stored on the computer-readable medium can produce an article of manufacture including instructions that implement the functions / acts specified in one or more blocks of the flowcharts and / or block diagrams.

[0071] Computer program instructions can also be loaded into a computer or other programmable data processing apparatus or other device to cause the computer or other programmable data processing apparatus or other device to perform a series of operational steps to create a computer-implemented process, whereby the instructions executed by the computer or other programmable device are referred to as a flowchart. The present invention provides a process for implementing the functions / acts specified in one or more blocks of the chart and / or block diagram.

[0072] Some embodiments of the present disclosure will now be described, by way of example only, with reference to the accompanying drawings. Specific reference will now be made in detail to the drawings, it being stressed that the particulars shown are by way of example and for purposes of illustrative discussion of embodiments of the present disclosure. In this regard, the description taken together with the drawings will make apparent to those skilled in the art how embodiments of the present disclosure may be practiced. [Brief explanation of the drawings]

[0073] [Figure 1A] 1 is a schematic diagram of an aircraft with a rotor-driven wing having a rotor oriented at a fixed skew angle relative to the horizontal flight pitch axis of the wing, in accordance with some embodiments of the present disclosure. [Figure 1B] FIG. 1 is a schematic diagram of a propeller-driven wing mounted to an aircraft fuselage with a rotor oriented at a fixed oblique angle relative to the wing's horizontal flight horizon, according to some embodiments of the present disclosure. [Figure 1C] 1A-1C are schematic diagrams illustrating airflow through a forward rotor, an aft rotor, and over a wing of an aircraft having rotors oriented at a fixed skew angle relative to the wing, respectively, in accordance with some embodiments of the present disclosure. [Figure 1D] FIG. 1 is a schematic diagram of an aircraft landing gear arrangement with wings having rotors oriented at a fixed skew angle relative to the horizontal flight pitch axis, according to some embodiments of the present disclosure. [Figure 1E] 1 is a schematic diagram of a rotor-driven aircraft with a wing having a rotor oriented at a fixed skew angle relative to the horizontal flight pitch axis of the wing, in accordance with some embodiments of the present disclosure. [Figure 1F] 1 is a schematic diagram of a rotor-driven aircraft with a wing having a rotor oriented at a fixed skew angle relative to the horizontal flight pitch axis of the wing, in accordance with some embodiments of the present disclosure. [Figure 1G] 1 is a schematic diagram of a rotor-driven aircraft with a wing having a rotor oriented at a fixed skew angle relative to the horizontal flight pitch axis of the wing, in accordance with some embodiments of the present disclosure. [Figure 1H]FIG. 1 is a schematic diagram of a rotor-driven aircraft with a wing and fuselage, the rotors oriented at a fixed skew angle relative to the horizontal flight pitch axis of the wing, in accordance with some embodiments of the present disclosure. [Figure 1I] FIG. 1 is a schematic diagram of a rotor-driven aircraft with a wing and fuselage, the rotors oriented at a fixed skew angle relative to the horizontal flight pitch axis of the wing, in accordance with some embodiments of the present disclosure. [Figure 1J] FIG. 1 is a schematic diagram of a rotor-driven aircraft with a wing and fuselage, the rotors oriented at a fixed skew angle relative to the horizontal flight pitch axis of the wing, in accordance with some embodiments of the present disclosure. [Figure 1K] FIG. 1 is a schematic diagram of a rotor-driven aircraft with a wing and fuselage, the rotors oriented at a fixed skew angle relative to the horizontal flight pitch axis of the wing, in accordance with some embodiments of the present disclosure. [Figure 1L] FIG. 1 is a schematic diagram of a rotor-driven aircraft with a wing and fuselage, the rotors oriented at a fixed skew angle relative to the horizontal flight pitch axis of the wing, in accordance with some embodiments of the present disclosure. [Figure 1M] FIG. 1 is a schematic diagram of a rotor-driven aircraft with a wing and fuselage, the rotors oriented at a fixed skew angle relative to the horizontal flight pitch axis of the wing, in accordance with some embodiments of the present disclosure. [Figure 1N] FIG. 1 is a schematic diagram of a rotor-driven aircraft with a wing and fuselage, the rotors oriented at a fixed skew angle relative to the horizontal flight pitch axis of the wing, in accordance with some embodiments of the present disclosure. [Figure 1O] FIG. 1 is a schematic diagram of a rotor-driven aircraft with a wing and fuselage, the rotors oriented at a fixed skew angle relative to the horizontal flight pitch axis of the wing, in accordance with some embodiments of the present disclosure. [Figure 1P] 1 is a schematic diagram of a rotor-driven aircraft with wings and a twin-fuselage airframe with rotors oriented at a fixed skew angle relative to the direction of the wings, according to some embodiments of the present disclosure. [Figure 1Q]1 is a schematic diagram of a rotor-driven aircraft with wings and a four-seat fuselage, with rotors oriented at a fixed skew angle relative to the direction of the wings, according to some embodiments of the present disclosure. [Figure 2] FIG. 1 is a schematic diagram of a rotor orientation transition mechanism for an aircraft according to some embodiments of the present disclosure. [Figure 3] FIG. 1 is a schematic diagram of a rotor orientation transition mechanism for an aircraft according to some embodiments of the present disclosure. [Figure 4] FIG. 1 is a schematic diagram of a rotor orientation transition mechanism for an aircraft according to some embodiments of the present disclosure. [Figure 5] FIG. 1 is a schematic diagram of a rotor orientation transition mechanism for an aircraft according to some embodiments of the present disclosure. [Figure 6] FIG. 1 is a schematic diagram of a rotor orientation transition mechanism for an aircraft according to some embodiments of the present disclosure. [Figure 7] FIG. 1 is a schematic diagram of an electric motor arrangement that can change the angle of attack of the propeller blades during propeller rotation by applying differential thrust between two parts of the electric motor, according to some embodiments of the present disclosure. [Figure 8] FIG. 1 is a schematic diagram of an electric motor arrangement that can change the angle of attack of the propeller blades during propeller rotation by applying differential thrust between two parts of the electric motor, according to some embodiments of the present disclosure. [Figure 9] FIG. 1 is a schematic diagram of an electric motor arrangement that can change the angle of attack of the propeller blades during propeller rotation by applying differential thrust between two parts of the electric motor, according to some embodiments of the present disclosure. [Figure 10A] 1 is a schematic diagram of aircraft attitude changes during takeoff and / or landing operations in accordance with some embodiments of the present disclosure. [Figure 10B] 1 is a schematic diagram of aircraft attitude changes during takeoff and / or landing operations in accordance with some embodiments of the present disclosure. [Figure 10C] 1 is a schematic diagram of aircraft attitude changes during takeoff and / or landing operations in accordance with some embodiments of the present disclosure. [Figure 10D] 1 is a schematic diagram of an aircraft having forward and aft rotors at different angles relative to one another, in accordance with some embodiments of the present disclosure. FIG. [Figure 10E] 1 is a schematic diagram of an aircraft having forward and aft rotors at different angles relative to one another, in accordance with some embodiments of the present disclosure. FIG. [Figure 10F] 1 is a schematic diagram of an aircraft having forward and aft rotors at different angles relative to one another, in accordance with some embodiments of the present disclosure. FIG. [Figure 10G] 1 is a schematic diagram of an aircraft having forward and aft rotors at different angles relative to one another, in accordance with some embodiments of the present disclosure. FIG. [Figure 10H] 1 is a schematic diagram of an aircraft having forward and aft rotors at different angles relative to one another, in accordance with some embodiments of the present disclosure. FIG. [Figure 10I] 1 is a schematic diagram of an aircraft having forward and aft rotors at different angles relative to one another, in accordance with some embodiments of the present disclosure. FIG. [Figure 11A] FIG. 1 is a schematic diagram of an aircraft configured for manned flight with wings having rotors oriented at a fixed skew angle relative to the horizontal flight pitch axis of the wing, in accordance with some embodiments of the present disclosure. [Figure 11B] FIG. 1 is a schematic diagram of an aircraft configured for manned flight with wings having rotors oriented at a fixed skew angle relative to the horizontal flight pitch axis of the wing, in accordance with some embodiments of the present disclosure. [Figure 11C] FIG. 1 is a schematic diagram of an aircraft configured for manned flight with wings having rotors oriented at a fixed skew angle relative to the horizontal flight pitch axis of the wing, in accordance with some embodiments of the present disclosure. [Figure 11D] FIG. 1 is a schematic diagram of an aircraft configured for manned flight with wings having rotors oriented at a fixed skew angle relative to the horizontal flight pitch axis of the wing, in accordance with some embodiments of the present disclosure. [Figure 11E] FIG. 1 is a schematic diagram of an aircraft configured for manned flight with wings having rotors oriented at a fixed skew angle relative to the horizontal flight pitch axis of the wing, in accordance with some embodiments of the present disclosure. [Figure 12A] 1 is a flowchart that schematically illustrates a method for reorienting an aircraft during takeoff of the aircraft, in accordance with some embodiments of the present disclosure. [Figure 12B] 1 is a flowchart that schematically illustrates a method for reorienting an aircraft upon landing of the aircraft, in accordance with some embodiments of the present disclosure. [Figure 13] FIG. 1 is a schematic block diagram of an electric motor (per se), according to some embodiments of the present invention. [Figure 14A] 1 is a schematic diagram of a tilting propeller guard for an aircraft, according to some embodiments of the present disclosure. [Figure 14B] 1 is a schematic diagram of a tilting propeller guard for an aircraft, according to some embodiments of the present disclosure. [Figure 14C] 1 is a schematic diagram of a tilting propeller guard for an aircraft, according to some embodiments of the present disclosure. [Figure 15] FIG. 1 is a schematic diagram of power connections to multiple battery units connected to multiple motors, according to some embodiments of the present disclosure. [Figure 16] FIG. 1 is a schematic diagram of an aircraft, optionally configured for manned flight, with a wing having a rotor oriented at a fixed skew angle relative to the horizontal flight pitch axis direction of the wing and an auxiliary rotor, in accordance with some embodiments of the present disclosure. [Figure 17] FIG. 1 is a schematic diagram of a control unit of a distributed flight control system, according to some embodiments of the present invention. DETAILED DESCRIPTION OF THE INVENTION

[0074] The present invention relates to the field of airborne flying vehicles, and more particularly to electric aircraft with vertical take-off and landing as well as geostationary flight capabilities.

[0075] (Overview) A broad aspect of some embodiments of the present disclosure relates to an unmanned or manned winged aerial vehicle (aircraft) equipped with multiple powered propellers oriented at a fixed skew angle relative to the direction of at least one wing of the aircraft. This configuration feature is referred to herein as a "fixed wing-to-rotor skew angle," and further details and definitions are provided below. By way of example, the "skew angle" of the wing relative to the propeller is selected from any of the following ranges: between 5° and 45°; between 10° and 35°; between 15° and 30°; and between 20° and 25°.

[0076] In some embodiments, the aircraft is configured for at least one of vertical takeoff, oblique takeoff (considered herein as a type of vertical takeoff, as defined below), and short takeoff. In some embodiments, the aircraft is configured for at least one of vertical landing and short landing.

[0077] Possible advantages of vertical take-off and / or landing of aircraft include the construction of soil for use by the aircraft. These include reduced ground space requirements, reduced dedicated space required for takeoff and landing (e.g., parking lots can double as landing sites), and improved safety during takeoff and landing (e.g., lower speeds near the ground).Possible advantages of fixed-wing (non-rotor) aircraft include the efficient and reliable generation of lift from forward motion through the air alone.

[0078] Aircraft designs have sought to combine vertical take-off and landing (VTOL) capabilities with fixed (non-rotating) wing flight support. Note that the term "fixed wing" is used in the art to mean "a wing that does not rotate on its center for lift generation," as opposed to the rotors of, for example, a helicopter, multicopter, or autogyro, and this is the meaning of the term as used herein. The meaning of other uses of the term "fixed" is explained below.

[0079] There are several potential issues associated with and addressed by efforts to combine VTOL and fixed wings. Broadly, these issues include vertical / horizontal flight transition reliability, added mechanical complexity, added weight, and other disadvantages rooted in the engineering compromises required to achieve a hybrid of both vertical and forward flight with a non-rotorcraft aircraft.

[0080] Some designs simply provide separate vertical and horizontal thrusters (at the expense of increased weight). In response, dynamic and static transition solutions have been proposed and implemented to potentially avoid "wasting" the vertical thrust capability of fixed-wing VTOL aircraft during forward flight. Dynamic implementations allow thrust to be redirected relative to the aircraft's airframe, for example, by rotatable rotor mounts and / or the use of baffles. However, dynamic solutions involve increased mechanical complexity, particularly during flight transitions, and also involve potential safety issues. Addressing these issues also entails associated potential increases in control complexity and cost.

[0081] Static transition strategies such as tail-sitters have also been proposed and implemented, in which essentially the entire aircraft transitions 90° from a vertical to a horizontal orientation. The same propulsors are used throughout, initially vertically to provide lift for takeoff, and then gradually banked to add horizontal speed, increasing bank until sufficient lift is achieved to complete the 90° transition to the aircraft's cruise direction. A reverse transition can be made to allow landing, but with the added safety complication of canceling the aircraft's forward speed without missing the target landing zone.

[0082] As new technologies become available, solutions that were previously unavailable or unnoticed may become potentially viable. For example, with the rise of small yet powerful electric motors, complementary increases in electrical storage capabilities, and the advent of inexpensive, powerful control electronics, a variety of aircraft powered by multiple propeller-coupled electric motors (herein "rotors") have been developed, are in development, or are expected to be developed. This general area of ​​endeavor is referred to herein as "multi-rotor technology."

[0083] There has been some recognition that the lift generated by fixed wings (e.g., according to the design approach described above) can also be utilized in conjunction with multi-rotor technology.

[0084] Surprisingly, the inventors have realised that fixed (non-rotating) wings with fixed wing-to-rotor skew, optionally in combination with other design features described herein, offer potential advantages for at least one phase of forward flight (cruise), takeoff and landing (particularly, but not exclusively, when combined with multi-rotor technology). In particular, for aircraft of sufficient size to carry cargo and / or passengers, e.g., unladen mass of 50 kg or more and / or laden mass of 120 kg or more, optionally This potential advantage arises for aircraft including those capable of carrying multiple passengers, e.g., 2, 3, 4, 6, 8, 10 or more passengers. Particularly for aircraft in these size ranges or larger, considerations of safety, robustness, simplicity, reliability, range, efficiency, power reserves, and / or responsiveness to controls are critical, e.g., related not only to basic functionality but also to regulatory approval and / or market acceptance.

[0085] In some embodiments, in the forward flight mode of the aircraft, the lift generated by the drive propellers is directed both forward and upward, while the lift generated by the wings is parallel to and opposite the direction of gravity's pull. This mode of operation offers potential energy efficiency advantages over using drive propellers alone for lift generation, as long as the wing-generated lift is configured to contribute more to flight support than its weight and drag contribute to flight loads. Because the wings do not carry the entire weight of the aircraft (at least in standard operation), they can be optionally designed to be lighter, thinner, and / or smaller than an all-wing design, potentially increasing the wing's net contribution to flight support. Rotor operation and / or design also offers potential advantages, as long as the rotors can optionally operate at low thrust conditions, allowing the fixed wing's lift to reach forward speeds sufficient to contribute to flight support. Because the sustained rotor thrust output requirement is reduced, the weight, power, and / or speed of rotor components (e.g., electric motors and / or propellers) can optionally be reduced.

[0086] In some embodiments, the aircraft is configured with a multi-rotor design, and control (stabilization and steering) of the aircraft is achieved by applying differential thrust to the aircraft's rotors. Flight control software can reduce or eliminate the need for flight surface stabilization for at least one of yaw, roll, and pitch during forward flight. Reduced control surfaces potentially improve efficiency. Conversely, the flight control software is optionally configured to reduce (optionally completely eliminate) the exercise of stabilization control in one or more axes at higher speeds, as this is driven by the aircraft's aerodynamics. For example, an aircraft with an aft-mounted vertical stabilizer is potentially protected from yaw at sufficiently high forward speeds. Instantaneous yaw tends to orient the stabilizer's vertical plane in a direction that is pushed back by the airflow, which restores a forward-facing yaw direction. Similarly, aerodynamic pitch angle is optionally established by a balance of forces between the wing and horizontal stabilizer (or other multi-wing configuration). This tends to keep the aircraft's pitch within a sufficiently narrow range, allowing the flight software to abandon exercise of pitch control during forward flight. Activation of stability control de-escalation and de-escalation (i.e., de-escalation and de-escalation of flight software control) in one or more axes may be automatic, driven by measured forward speed, or may be manual.

[0087] The fixed wings optionally do not include control surfaces themselves, which has potential advantages for simplicity of design and / or weight, or they do include control surfaces, for example, for emergency control in the event of loss of power to one or more rotor motors and / or for added maneuverability.

[0088] However, the potential benefits are not limited to effects on forward flight characteristics, and different fixed blade-to-rotor splay configurations have different potential advantages.

[0089] The inventors have realized that some potential short-range aircraft applications have similar energy budget requirements for forward flight and hovering (e.g., energy budget requirements within about 1-2 times of each other). Relatively long-range, highly efficient forward flight is balanced by relatively short-range, but less efficient, hovering flight, especially when allocating a safety margin portion of the energy budget to hovering flight. This allocation is useful for determining the energy budget for a given point-to-point flight. This is a reasonable choice for forward flight, since the need for a safety margin may arise due to obstacles at the landing site that require loiter and / or hover maneuvers to avoid and / or wait until they clear. Thus, for example, a 30-minute forward flight capability may be provided with an assumed hover capability of approximately 4 minutes (including the safety margin). In some embodiments, hover flight uses approximately five times more power than forward flight (e.g., due to the lack of aerodynamic lift advantages), resulting in a nominal balance of energy budgets between forward and hover flight in a ratio of approximately 3:2.

[0090] An aspect of some embodiments of the present disclosure relates to forward flight of an aircraft equipped with a wing-to-rotor skew angle, where the rotor is mounted relative to the wing such that its thrust airflow is directed in a different direction than the airflow experienced by the wing in forward flight. In some embodiments, the wing-to-rotor skew angle is fixed.

[0091] In some embodiments, the rotors are mounted within the span of the wing (e.g., closer to the fuselage than the tip of the wing). In some embodiments, the mounting configuration is such that at least one rotor is mounted with its center below and forward of the wing and at least one rotor is mounted with its center above and aft of the wing. In this configuration, optionally, the rotors are mounted in a generally planar configuration and are tilted (about the pitch axis) relative to the orientation of the wing.

[0092] In some embodiments of the present disclosure, during forward cruise flight (e.g., at speeds greater than about 55 km / h), both the wing and the propeller generate lift, e.g., the wing generates at least 10% of the lift, 25% of the lift, 30% of the lift, 50% of the lift, "the majority of the lift" (more than 50% of the lift), or 70% of the lift, and the propeller provides the remaining lift, comprising at least 10% lift. From these examples, it can be seen that it is a potential advantage that the lift generated by the wing is a relatively large portion of the total lift (e.g., the majority of the lift) for forward flight, and that using wing lift optionally reduces the need to use the vertical component of rotor thrust to stay airborne, with the total vertical and horizontal thrust required for forward flight typically being less than the vertical thrust required for horizontally stationary hovering flight.

[0093] To the extent that both rotors and wings are used to generate lift, there is a potential advantage in arranging and configuring them to avoid interfering with each other's lift-generating functions. Both propellers and wings are intended to operate in a laminar flow of air (at least upon initial impact), and when turbulence occurs, they tend to introduce vibrations and reduced efficiency that are important to aircraft performance. The initial airflow (wash) from the propeller is turbulent and is directed in the same direction as the thrust. Wings also leave turbulence in their wake. Therefore, when both wings and propellers are operating, there is a potential benefit in moving their "output" (wake / airflow) away from each other's "input" (leading edge / intake). For purposes of the description provided herein, a blade is said to avoid interference with rotor thrust as long as it does not intersect a right circular cylinder centered on the rotor, i.e., a right circular cylinder having a central longitudinal axis coincident with the axis of thrust generated by the rotor (or other type of motor assembly) and having a radius equal to the radius of the rotor's propeller, turbine blades, jet exit opening, or other primary flow generating / flow shaping element. If the blade intersects the air turbulence induced by the motor assembly outside of this cylinder, it is considered, at least for purposes of this description, to have no such interference with the motor assembly thrust itself.

[0094] The lateral position of the rotor center (i.e., the distance along the pitch axis from the aircraft center) optionally ranges from the aircraft center position to outside the aircraft's wingspan. Rotors not at the aircraft center are aligned with a corresponding pair of rotors to maintain balance and equalize flight stresses. Preferably, the rotors are provided as a single rotor (one on each side). Potentially, a multi-rotor aircraft (e.g., a multicopter with six or more rotors) could continue to fly safely (optionally with appropriate adjustments to the controls) even if one or more rotors are inoperable or missing.

[0095] A particular advantage of mounting the motors within the wing span is reduced stress on the mounts. The mounting members are optionally shorter, for example, reducing the leverage forces at the joints that attach them to the rest of the aircraft. Avoiding mounting the motors at or near the tips of the wing (as another example) potentially allows the wing itself to be constructed lighter, potentially resulting in a reduction in empty weight. Flight stability is potentially improved by mounting the rotors farther apart, but this also tends to increase the leverage stresses on the mounts. By mounting the motors to the front and rear of the fuselage, the mounting members of some embodiments are kept even shorter, taking advantage of the stiffness of the fuselage itself.

[0096] Furthermore, as long as the wings and rotors are independent of each other in terms of their placement, the wings can be designed to be long and / or thin to suit their function of providing lift, without concern for the rotor extending into the airstream or for stiffening the rotor itself to support and / or transmit rotor thrust to the aircraft.

[0097] The concept of mounting motors both forward and aft on the wings allows the aircraft to operate (at least in some flight conditions) with the flight characteristics of a wingless multirotor aircraft, with the center of gravity supported between the rotors (including between them in a plane perpendicular to the pull of gravity). In the same configuration, the wings can also exert their lift at or near the center of gravity (at least in high speed forward flight).

[0098] An aspect of some embodiments of the present disclosure relates to takeoff and / or landing strategies for aircraft equipped with fixed wing-to-rotor splay angles, which in particular offer potential advantages over a range of such strategies, assisted, for example, by the provision of appropriately configured landing gear.

[0099] In a vertical takeoff mode of an aircraft, in some embodiments, the propellers are oriented (e.g., by pitch axis rotation of the entire aircraft) to exert a lift force parallel to and opposite to the pull of gravity (and concomitantly tilting the wings from their horizontal forward flight orientation according to a fixed cant angle). This mode can be used to climb or descend the aircraft in a hover mode, including taking off or landing without taxiing. A vertical takeoff is without taxiing, or without fixed-wing lift generation, or both. Optionally, a vertical takeoff is preceded by a maneuver that levels the initially tilted propellers by lifting one side of the aircraft while keeping the other side in contact with the ground.

[0100] In an aircraft's oblique takeoff mode, the aircraft launches at an angle relative to the ground (without taxiing or significant prior fixed-wing lift, or both). Optionally, the propeller is originally oriented at an oblique angle relative to the ground ("tilted"), and the direction of oblique takeoff is perpendicular to the direction of the propeller. Optionally, the wings are pitched during oblique takeoff in a direction for horizontal forward flight. However, there is no restriction on oblique takeoff that precludes other wing pitch tilts, such as forward tilt (the aircraft's nose pitches further downward) or rearward tilt (the aircraft's nose pitches further upward).

[0101] During angled takeoff, the angled propeller generates both forward thrust and lift thrust. The transition to thrust sufficient for takeoff is rapid enough that the forward thrust does not result in forward running on the ground. Or there is still only weak forward acceleration at takeoff speed, with the wings contributing less than 10% to the lift acting on the aircraft. If forward travel occurs, it may be an accidental movement where ground friction is overcome during thrust build-up. Optionally, the forward movement in an oblique takeoff extends over a distance less than the length of the aircraft.

[0102] After the oblique takeoff, the aircraft may optionally be reoriented in the air to a vertical thrust (horizontal propeller) mode, which may or may not cancel the forward velocity generated during the initial takeoff. Insofar as the characteristic that distinguishes vertical takeoff from other modes of aircraft launch is its lack of reliance on fixed-wing lift generation for forward motion, oblique takeoff can be considered (and is generally treated as such herein) to fall within the vertical takeoff category. Thus, "vertical takeoff" includes oblique takeoff herein, unless the description specifically relates to a static hover.

[0103] Landing involves the reverse of the takeoff orientation, although maneuvers to reduce or cancel ground speed may be included. If not fully canceled before touchdown, it results in an "inclined" landing, in which the aircraft slows at least partially on the ground. As used herein, a "vertical landing" includes landing from a hover without significant fixed-wing lift being generated (e.g., less than 10% of total lift), even if there is a non-zero ground speed at touchdown. However, the embodiments described herein may envision entering a ground-static hover to avoid any indication of "instant braking" before landing on a small object, such as a landing pad. After the initial touchdown (e.g., touching the ground while in a ground-static hover), the aircraft may reorient and return its rotors to an inclined position upon further rest.

[0104] Vertical / diagonal landings and takeoffs are encompassed herein by the abbreviation VTOL (vertical takeoff and / or landing; note that this abbreviation is often defined with the conjunction "and"). It should be understood that the abbreviation VTOL is intended to describe the capabilities of an aircraft and does not limit its functionality to vertical takeoff and landing alone. Additionally, the abbreviation VTOL should not be understood to assert that vertical takeoff and vertical landing are necessarily coupled.

[0105] As an example of non-vertical takeoff and landing operation, in a short takeoff mode of some embodiments of the aircraft described herein, the aircraft taxis a short distance over the ground before takeoff. Here, "short takeoff mode" is fixed-wing assisted, and the airspeed at takeoff is sufficient for the wings to contribute at least 10% of the total lift. In some embodiments, the aircraft is capable of a short landing, in which the aircraft touches down in forward motion with at least 10% of the residual lift generated by the fixed wing. By definition, a short takeoff allows the aircraft to clear a 15-meter obstacle within 450 meters of the start of the takeoff run.

[0106] One aspect of some embodiments of the present disclosure relates to a winged aircraft configured with powered propellers oriented at a fixed oblique angle relative to the wings, such that, upon landing, the propellers are also held at an oblique angle relative to the ground.

[0107] The basic "tailsitter" design mentioned briefly above can rotate the position of the occupants and / or cargo by 90 degrees, which can be uncomfortable and can result in significant cargo shifting. Mitigating such extreme changes in angle, for example, with rotating seats or compartments, potentially requires overcoming fuselage, cabin, and / or cockpit design challenges.

[0108] In some embodiments of the present disclosure, such large rotations are avoided from the start. The orientation adjustment from forward flight to hovering flight is, for example, the same as the fixed wing-to-rotor skew angle, preferably less than 45°, for example, in the range of about 15° to 30°, and preferably less than about 20°. ~25°. In some embodiments, tilting the seats or compartments by a few degrees may further reduce the effective angle change experienced by cargo on the aircraft.

[0109] While the continued reliance on rotor thrust for lift during forward flight means that lift efficiency is potentially lower than with pure fixed-wing lift, there is still a potential advantage in lift efficiency gains relative to the baseline of a wingless multirotor-powered aircraft.

[0110] Additionally, there are potential advantages when the aircraft cargo experiences different orientations relative to the ground.

[0111] The inventors have realized that a surprising alternative to a two-state transition is a three-state transition, which is optionally implemented using a rotor-driven aircraft with wings and rotors oriented at oblique angles to one another. Alternatively, in some embodiments, directional transitions can optionally be avoided for takeoff and / or landing.

[0112] In some embodiments, upon takeoff or shortly thereafter, the aircraft first transitions from a landing condition in which the rotors (and at least one of the net combined rotor thrust vector) are tilted away from vertical to an intermediate vertical takeoff condition (optionally starting with the landing gear also touching the ground) in which the rotors and at least one of the net combined rotor thrust vector are vertical. Then, after takeoff and / or after sufficient altitude is attained, the aircraft transitions to a third condition in which the rotors and at least one of the net combined rotor thrust vector are again tilted away from vertical, optionally in the same or a different direction than that of the landing condition. Then, in some embodiments, seats or other onboard equipment are adapted to be most comfortable for the aircraft's landing configuration. This orientation may be disturbed only during takeoff, after which the comfortable orientation is restored during normal flight.

[0113] In oblique takeoff mode, a variation of vertical takeoff, no intermediate turbulence is required, and the same orientation is optionally maintained throughout the transition from the aircraft on the ground to full forward flight, which has the potential advantage of reducing the maximum tilt angle of the aircraft cargo during takeoff.

[0114] For landing, the procedure is reversed: the transition from a hovering state (e.g., rotors oriented vertically) to a landed state (e.g., rotors tilted away from vertical) optionally occurs in the air (possibly applying a brief horizontal acceleration before the aircraft touches down to allow braking on the ground) or after contact with the ground.

[0115] As previously mentioned, the seating position of the occupants on the ground is optionally the same both before takeoff and during flight, with the "tilted seating" phase being limited to vertical takeoff, landing, and possibly the period of speed reduction before landing. Optionally, some difference in seating orientation remains in each case. For example, the aircraft in forward flight is in the "most comfortable" orientation for those seated inside, with the rotors tilted forward by a fixed wing-to-rotor skew angle. On the ground, the aircraft tilts slightly backward from this angle (e.g., within a range of approximately 1° to 10°) to provide the occupants with a moderate reclining sensation. It should be understood that a moderate range of chair adjustment angles (e.g., up to approximately 10° to 15°) can be used to further alter the seating angle of the chair as deemed appropriate.

[0116] A tilt rotor landing condition is achieved in some embodiments by appropriately selecting the relative lengths of the rotor attachment arms and landing gear. Placing the forward mounted rotor under the wing potentially has the effect of placing it very close to the ground, reducing the rotor's ability to catch dust and debris. Note in particular the potential susceptibility to interaction with debris. Optionally, longer landing gear and / or a moderate "tilt backward" is used to mitigate this weakness. Optionally, the landing gear itself is adjustable to change the orientation of the aircraft on the ground (e.g., tilt forward for loading cargo and / or boarding passengers, tilt backward before takeoff to keep the rotors from contacting the ground, etc.).

[0117] In some embodiments, slowing the aircraft includes reducing the forward pitch of the aircraft. In this configuration, the wings potentially act as brakes, at least at high speeds. Optionally, the pitch of the aircraft's rotors is reversed (past a vertical net thrust vector position) to impose further slowing. Optionally, the aircraft is controlled to yaw through about half a rotation while maintaining at least a portion of its forward pitched rotor pitch, and then slowed for a period of time while flying backward through the air. Optionally, such a maneuver is performed after the airspeed has dropped sufficiently so that the wings are no longer causing the aircraft to perturb beyond the compensation capabilities of the control system.

[0118] Optionally (e.g., under autopilot and / or auto-assisted piloting), horizontal deceleration (optionally during reverse flight) reduces ground speed to zero approximately simultaneously with touchdown of the aircraft. A potential benefit of this is to keep the aircraft more level during descent, thereby increasing crew comfort and reducing cargo shifting and the risk of the aircraft becoming unbalanced.

[0119] In some embodiments, (optionally relatively low-power) auxiliary motors are provided that operate (particularly during landing and / or takeoff) to provide additional thrust as needed for hovering, horizontal maneuvers, acceleration, deceleration, and / or balancing opposing thrust. This has the potential advantage of allowing low acceleration maneuvers with fewer bank adjustments (e.g., relative to pitch). The auxiliary motors are optionally lightweight to reduce weight penalty. Optionally, they are mounted to feather, retract, and / or pivot the blades out of the airflow for drag reduction during high-speed forward flight. Optionally, they are operated to generate additional thrust during forward flight, and optionally used in reverse, for example, during landing maneuvers. In some embodiments, auxiliary engines are convenient but not necessary to operate the aircraft or otherwise maintain airworthiness. In some embodiments, the auxiliary engines provide redundancy, for example, for stabilization. For example, an event such as a collision (e.g., a collision with a tall obstacle that damages the motor mounting arms or a motor mounting arm carrying the motors) could potentially disable the main motors and create flight instability that cannot be fully compensated for by the remaining motors. In such a condition, the auxiliary motors can be activated to help maintain at least some of the flight stability, for example, to balance the aircraft long enough to perform an emergency landing.

[0120] For example, in an eight-motor (or other) coaxial arrangement with two motors on each arm, if an external impact from a pole were to disable two motors on the same arm, for example, the vehicle could lose stability and tip over. The additional two motors would maintain stability in such a case.

[0121] An aspect of some embodiments of the present disclosure relates to a self-tilting rotor mounting assembly for an aircraft, comprising a pivot bar and a plurality of rotors mounted on either side of the pivot bar and pivot mounts (also referred to herein as hinges) of the pivot bar, the rotors operable to rotate the bar about the pivot mounts and pitch axis.

[0122] In some embodiments, the pivot angle range of the pivot bar is limited between a relative horizontal orientation and a relative vertical orientation. The orientation is selected by varying the relative power to the rotors on either side of the pivot mount. The result is a bistable mechanism that can be selectively converted between two modes by varying the relative power to the rotors. In one mode, the pivot bar is fixed to the more horizontal side of the pivot angle range, and in the other mode, the pivot bar is fixed to the more vertical side of the pivot angle range.

[0123] In some embodiments, mode switching control is performed in part using forces acting on the aircraft's center of gravity and center of lift. In some embodiments, the aircraft's center of gravity and its aerodynamic center of lift are both located forward of the pivot center of the pivot mount. The pivot angle range of the pivot bar is limited so that in hover flight, the weight of the aircraft locks the fuselage and pivot bar in a relative orientation appropriate for generating hover thrust. As lift generated in forward flight releases weight-induced forces from the pivot, the pivot bar is forced to rotate more vertically (e.g., by appropriate adjustment of the relative power of the forward and aft rotors), while the rest of the aircraft's pitch remains stabilized by aerodynamic forces. Optionally, the pivot angle range is also limited so that appropriate balance of rotor thrust "locks" the mounting assembly into a fixed forward flight configuration.

[0124] Optionally, the pivot bars are located on both the left and right sides of the aircraft, and the pivot bars are coupled together to form a single pivot frame. Optionally, the pivot bars are not mechanically connected, but cooperate, for example, by flight control software. Optionally, one pivot bar is held in a position (e.g., at the rear of the aircraft) that allows the rotor to move freely without hitting the body of the aircraft.

[0125] In some embodiments, the pivot mount is positioned so that its pivot center is offset longitudinally and forward of the center of gravity of the tilt rotor mounting assembly. With the motors off, the tilt rotor mounting assembly therefore tends to rotate clockwise under gravity when viewed from the side with the front of the aircraft to the left. This tendency to rotate is mechanically stopped at a position that keeps the rotor in the proper orientation for hover flight, e.g., when the rotor starts, initially with the rotor thrust vector pointing vertically downward. 2. Once airborne, the pivot mount's attitude is no longer controlled by the pivot mount's center of gravity, but instead is controlled by differential thrust forces.

[0126] In some embodiments, the center of gravity of the aircraft fuselage and wings is located forward of the pivot center. As a result, when the aircraft climbs under rotor thrust, the aircraft fuselage and wings tend to rotate in a pitch down direction. This pitch down movement also tends to pull the pivot mount in a pitch down direction because the pivot mount cannot move any further in a relative pitch up direction compared to the fuselage. As a result, the self-tilting rotor mounting assembly and fuselage-wing assembly remain fixed to one another substantially as they would be on the ground, even though the force sources are different.

[0127] Overall aircraft pitch can be controlled by providing relatively different thrust to the forward rotor compared to the aft rotor, including thrust that balances any tendency of the center of gravity to pitch the entire aircraft downward.

[0128] At this point, the rotors are reoriented to provide forward thrust. As speed increases, the wings begin to generate forward lift in response, from their lift centers, which are also forward of the pivot center (and in some embodiments, forward of the wing and fuselage center of gravity). This lift releases the "lock" that was placed on the pivot mount by the aircraft's weight. From this point on, thrust, which is relatively more distributed to the aft rotor than to the forward rotor, follows the aircraft fuselage and wing. This tends to rotate the pivot mount more vertically without the need for a vertical pulley, as the aircraft fuselage and wing are already held in a relative upward pitch by the lift forces of the wing.

[0129] Optionally, the rate of rotation transition is limited by pivot resistance, e.g., frictional resistance to the turn (e.g., skin friction and / or viscosity thereon) and / or inertial resistance (e.g., due to the flywheel), slowing the pitch rotational movement of the pivot mount (e.g., resulting in a gradual reorientation). In preparation for vertical landing, the upper (rear) rotor optionally provides a small amount of power relative to the lower (forward) rotor, returning them to a hover position. As airspeed decreases, the weight of the aircraft returns to being supported by the pivot mount, restoring hover lock.

[0130] One aspect of some embodiments of the present disclosure relates to a variable blade pitch rotor with two coaxial electric motors. In some embodiments, each of the two coaxial electric motors is hingedly attached to a different portion of the base of the rotor's propeller blades. Changes in the relative position of the electric rotor sections of each electric motor translate into changes in the pitch of the propeller blades. In some embodiments, only one electric motor is attached directly to the blades. Blade rotation is driven by a gear (e.g., ratchet and pinion) mechanism coupled to the relative position of the electric rotor section of the other electric motor. In some embodiments, blade pitch is variable across its entire range (from minimum pitch to maximum pitch) in less than one second. The range of pitch change is, for example, about 5°, 10°, 15°, or other ranges.

[0131] An aspect of some embodiments of the present disclosure relates to a propeller guard shaped to reduce aerodynamic drag when oriented obliquely relative to the propeller it surrounds. In some embodiments, the walls of the propeller guard are generally obliquely cylindrical, with the obliqueness of the cylinder selected so that each radial cross-section of the propeller guard wall points in the direction of horizontal forward flight of the aircraft. In some embodiments, at least a portion of the walls of the propeller guard are airfoil-shaped, optionally shaped to generate lift.

[0132] One aspect of some embodiments of the present disclosure relates to an electric motor comprised of a plurality of individually wound coils, each of which occupies a single position around the stator of the electric motor (e.g., is not separated from itself by the interposition of another coil). In some embodiments, each coil acts as or is associated with an electric and / or magnetic force sensor that detects the angular position of a rotor magnet. For each individual coil, a respective individual coil controller selects the polarity and / or magnitude of the current supplied to the coil based on the detection.

[0133] Optionally, the multiple individual winding coils are configured to cooperate to act as a single-phase electric motor (e.g., generate electromotive forces of substantially the same level and polarity at the same time) even though the coils are not electrically interconnected. Optionally, the individual controllers cooperate to act as a single-phase electric motor even though they are not coordinated with each other.

[0134] Optionally, all of the individually wound coils operate simultaneously to exert a force on the rotor of the electric motor. This is a potential advantage as it allows the electric motor to produce greater peak forces with a given number of coils. While the number of parts may potentially increase, the parts themselves may be lighter than would otherwise be required, for example, due to reduced power handling requirements for the individual coils. Another potential advantage is that the loss of a single coil (e.g., due to a short circuit) does not necessarily cause the electric motor to cease operation. Another potential advantage is that the coils themselves may be electrically connected to one another. The coils may be individually replaceable within the electric motor, for example as part of maintenance of the electric motor.

[0135] One aspect of some embodiments of the present disclosure relates to a battery unit power distribution arrangement in which each of multiple battery units can power a different corresponding set of rotors, where each set of rotors is comprised of rotors located at diagonally opposite corners of the aircraft. In some embodiments, this configuration acts to potentially prevent aircraft instability resulting from the loss of power from a single battery unit, which would cause a sudden imbalance in diagonally opposite thrust.

[0136] In some embodiments, a rotor having multiple propellers and a corresponding multiple power units for driving those propellers are optionally powered by a different battery unit for each power unit. Thus, in some embodiments, there are two battery units configured to power each rotor, and each battery unit also powers the power unit of the diagonally opposite rotor. This is also referred to herein as a "diagonally wired" power configuration.

[0137] The rotors are "diagonally opposed," so that any thrust imbalance delivered between diagonally opposed rotors will tend to rotate the aircraft about a diagonal axis of rotation approximately perpendicular to the diagonal on which the opposed rotors are mounted. Because the diagonal axis of rotation passes in close proximity to the other rotor, providing corrective stabilizing thrust requires providing a change in thrust of at least twice the imbalance thrust difference, and optionally more, e.g., five, ten, or more. Note that diagonally opposed rotors are also rotors in a "balancing" rotor set, as described in connection with FIG. 15 herein.

[0138] Thus, for example, if one power unit (e.g., one of multiple power units) of a diagonally opposed rotor pair fails completely, potentially no other rotors in that pair can alter their own thrust in time to overcome the thrust imbalance (which then propagates to the diagonally opposed rotor pair). In some embodiments (e.g., in a configuration with a rotor at each of the four corners of the aircraft), the diagonal axis of rotation resulting from a rotor failure could potentially pass so close to a subset of the remaining rotors that none of them could even operate to provide a sufficiently controlled counter-thrust to maintain aircraft stability. As a result, these remaining subsets of rotors are ineffective in providing recovery from even relatively small and / or transient imbalances. For example, in some embodiments, if one subset of rotors in a pair experiences a severe power failure (e.g., loses half or all of its power), rotation about the diagonal axis is initiated. When a rotation is detected (e.g., by an inertial measurement unit connected to the aircraft's control system), in some embodiments, the normal aircraft response is to reduce power on the other (diagonally opposite) subset of rotors. However, the rotation may already have reached a certain magnitude of momentum before it is detected, and either way, power control may not be able to immediately reduce power, or attempts to restore stability may result in oscillations in the aircraft's attitude, or both. This is potentially a particular problem for heavier aircraft with greater associated momentum. Even if this situation is not necessarily irreversible (e.g., the aircraft's own center of gravity may be low enough to help prevent flipping), it is potentially advantageous to eliminate or reduce the risk of control delays that could cause rotation about a diagonal axis and / or accumulation of rotational momentum, at least in some failure modes.

[0139] In a diagonally wired power configuration, a failure mode involving loss of battery unit power (e.g., loss of connection or malfunction of the battery unit itself) is potentially self-balancing in that the loss of power occurs essentially simultaneously between diagonally opposed motors, thus initiating thrust balancing as an inherent part of the failure itself, independent of the flight controller detecting the onset of rotation and / or commanding a correction.

[0140] In some embodiments, loss of power to a rotor power unit (e.g., due to a battery unit failure) is detected by the rotor controller. This is optionally used as a signal to automatically initiate an increase in power to another rotor power unit (e.g., even before a change in aircraft attitude is detected). This is a potential benefit, for example, to reduce control delays in response to a power loss failure mode. Optionally, the signaled power portion increase is temporary, e.g., the effect of the signal is reduced over time, e.g., in favor of attitude sensing-based control. Being temporary has the potential advantage of allowing a more immediate response (e.g., to reduce loss of flight equilibrium after a failure) without introducing permanent new factors that may themselves interfere with re-establishing flight equilibrium.

[0141] Optionally, the signal is propagated for use in controlling other rotors, for example to command or confirm (selectively, temporarily) the removal or reduction of power to the power units of diagonally opposite rotors, which has the potential advantage of reducing control delays in response to, for example, damage affecting only the power connections of one rotor.

[0142] One aspect of some embodiments of the present disclosure relates to an arrangement of multiple flight controller units, with each of the multiple rotors under the control of its own corresponding flight controller. Optionally, each flight controller is mounted at the location of the rotor it controls and, optionally, includes its own inertial measurement unit (IMU). This offers a potential advantage in terms of robustness, insofar as failure of one flight controller only affects the control of one flight controller. This also offers a potential advantage in reducing flight control resonances, which may occur, for example, due to strut and / or airframe flexing, because each rotor is under the control of a flight controller with an IMU, which measures the same local variations in rotor attitude due to airframe flexing.

[0143] In some embodiments, flight data (e.g., IMU data) is shared between flight controller units, with each flight controller having access to the same flight data describing the current flight state of the aircraft. Thus, each flight controller unit has access to the same description of the flight state of the entire aircraft. Optionally, data from each IMU of each flight controller unit is combined into a single estimate of the flight state of the aircraft using the same combination method for each flight controller unit (e.g., including the same algorithm and, optionally, at least one of weighting for averaging and the same method of rejecting exceptions, such as outlier data).

[0144] However, optionally, each flight controller unit has particular discretion over flight data measured by its own IMU in generating control commands to its particular rotor and / or evaluating the results of these control commands. In some embodiments, the flight controller optionally issues control commands directed at the overall flight state of the aircraft (e.g., calculated from shared flight data, optionally the same for all flight controller units). However, the flight controller optionally assigns a particular weighting to flight data from its own IMU to determine the rapid execution of commands, such as the rate of attitude change, which is dependent on the flight state of the aircraft. may be partially absorbed by local airframe flex before affecting an overall change in . This determination optionally influences how the control outputs of the flight controller are adjusted, potentially helping to dampen or prevent control oscillations.

[0145] An aspect of some embodiments of the present disclosure relates to yaw control for a multi-motor aircraft, where the rotor comprises multiple coaxially mounted counter-rotating propellers.

[0146] In some embodiments of the present disclosure, yaw control is achieved through at least one of a reaction to the change in rotational momentum as the rotor blades accelerate / decelerate and a rotation imparted to the air mass due to drag from the rotating rotor, referred to herein as "yaw torque." Additionally or alternatively, the rotor is oriented at a tilt about the roll axis (and optionally at least one axis perpendicular to an axis extending approximately between the rotor and the center of gravity of the aircraft). The tilt is a few degrees (e.g., 5° or less) inward or outward relative to the aircraft's median axis, which is parallel to the roll axis. This potentially provides the rotor with a thrust component that can be used to assist in yaw control, referred to herein as "yaw thrust." Preferably, the tilt distributes the thrust component in the yaw direction, approximately equal to or less than the rotor's ability to compensate for the use of yaw torque.

[0147] Optionally, in the case of a single-propeller rotor design, the tilt direction is chosen so that yaw thrust is applied to yaw torque. For example, increasing the speed of a rotor that rotates clockwise (as viewed from above) will impart yaw torque in the counterclockwise direction. The same rotor is also oriented so that increasing its thrust will result in counterclockwise yaw.

[0148] It is generally desirable to be able to perform yaw control without affecting forces applied in other directions (e.g., without causing changes in pitch, roll, or altitude). Thus, rotors are generally controlled so that rotors in opposing positions (e.g., diagonally opposite each other) simultaneously increase rotation / thrust in a ratio that avoids rolling or pitching the aircraft. To avoid changes in altitude, additional thrust from a rotor providing increased yaw torque / thrust is balanced by decreased thrust from the other rotor, typically the rotor providing yaw torque / thrust in the opposite direction, which decrease is itself preferably diagonally balanced. Thus, yaw control in (say) a clockwise direction is optionally performed, e.g., by increasing all counterclockwise rotating rotors and decreasing clockwise rotating rotors.

[0149] In some embodiments, the rotor is configured with multiple coaxially arranged propellers, e.g., two counter-rotating coaxial propellers per rotor. Yaw control using yaw thrust can be provided with such an arrangement, but if both coaxial propellers rotate faster at once, they may counteract each other's yaw torque effects and reduce available yaw control. Conversely, if (for example) clockwise propellers rotate faster and counterclockwise propellers rotate slower, the net change in yaw thrust may be reduced and yaw control using yaw thrust may be impaired. Particularly for heavier aircraft and / or with limited energy budgets, loss of yaw authority for both effects can be a significant disadvantage. Reducing available yaw control power impairs the aircraft's ability to control yaw direction, for example, against wind forces, even when the aircraft is grounded, which is a potential problem for safety and / or low-speed maneuverability. Even in still air, reduced yaw controllability can particularly increase the duration of the hover phase of flight, which is both (1) a phase of flight in which energy is used very rapidly and (2) a phase of flight that is likely to be considered to have allocated flight energy reserves, and therefore also deductions from the aircraft's rated range.

[0150] In some embodiments of the present disclosure, the flight controller is configured (at least under normal circumstances) to apply yaw torque to only one of the two propellers in each rotor with coaxially arranged propellers, and each rotor is controlled to increase or decrease both yaw thrust and yaw torque together. Thus, if a rotor is mounted at an angle that provides clockwise yaw thrust, its counterclockwise-rotating propeller is rotated faster to simultaneously apply both clockwise yaw torque and counterclockwise yaw thrust. The speed of the clockwise-rotating propeller is left unchanged because a faster rotation opposes the yaw authority applied via yaw torque and a slower rotation opposes the yaw authority applied via yaw thrust.

[0151] A mating rotor mounted diagonally (for example) to the rotor is preferably controlled in a similar manner to maintain net force so that pitch / roll is not disturbed. Since altitude is also maintained, a rotor tilted to provide counterclockwise yaw thrust preferably reduces net thrust by specifically reducing rotation of the clockwise rotating propeller, but not the counterclockwise rotating propeller. This has the additional effect of increasing net thrust about the yaw axis controlled in that direction.

[0152] In some embodiments, the controlled propellers are all upper propellers or all lower propellers of each rotor. For example, in a rotor configuration with an even number of rotors, each rotor with an upper propeller rotating clockwise and a lower propeller rotating counterclockwise is adjacent to two rotors with the opposite configuration (and vice versa). Each rotor is in turn adjacent to two rotors tilted to exert yaw thrust in the opposite direction to itself. For each rotor, the propeller used for control rotates in the opposite direction to the direction of application of yaw thrust for that rotor. One or more rotors may be excluded from yaw control, such as in the case of an odd number of rotors.

[0153] Optionally, for example by tilting motors on opposite sides of the aircraft in opposite directions, the opposite rotors reverse the relative top-to-bottom order of the clockwise and counterclockwise rotating propellers (and the top-to-bottom order is optionally reversed both right / left and front / rear).

[0154] In some embodiments, the flight controls optionally cease actively exercising yaw authority during forward flight above a certain airspeed except upon receiving a pilot command, which is possible for embodiments having sufficient vertically protruding aerodynamic surfaces to induce inherent yaw stability in the aircraft.

[0155] (definition) Forward flight motion in a plane perpendicular to gravity is referred to herein as "horizontal forward flight" or "parallel to the ground" flight. When "flight" is referred to without other qualification, it should be understood to be "horizontal forward flight."

[0156] An aircraft wing is said to be horizontal or parallel to the "ground," "plane of the Earth," or "surface" when the aircraft is in horizontal forward flight, regardless of the wing's actual shape or how its cross section is oriented. For example, an upturned wing is also said to be "parallel to the ground" during the wing's horizontal forward flight. Also, a "horizontal wing" (in the "horizontal flight pitch direction") depends on the direction of airflow at a given altitude, and is independent of the specifics of the wing chord direction and / or angle of attack within the context of forward flight. As used herein, a wing in horizontal flight and / or horizontal wing forward flight refers to the ideal situation in which the wing's direction is parallel to the flat ground and lift is in balance with gravity. At a constant speed, drag balances horizontal thrust in the forward flight direction. This may not necessarily be the case in forward flight, depending on the phase of flight and flight conditions such as aircraft loading. Although not a true condition of flight, it serves as a benchmark.

[0157] In contrast, "wing direction" should be understood to specifically refer to the chord pitch direction measured between the leading edge and the leading edge of the wing (the "pitch direction" of the wing refers to the same direction). Because the chord may be angled slightly differently in different cross sections, "wing direction" more specifically refers to the chord length-weighted average chord direction. In embodiments with multiple wings, the wing span- and chord length-weighted average chord direction of all wings is calculated. In cases where the chord direction is variable (e.g., its shape is variable), the chord intended here more specifically refers to the chord applicable to the wing configuration during most efficient forward flight and simultaneously at most efficient cruise speed.

[0158] Generally, wings are oriented to produce the ideal (most efficient) angle of attack during forward flight, which is approximately within 0° to 12° of the chord direction. The ideal angle of attack is the most efficient ratio of lift to drag across the wing while moving in still air at the aircraft's rated cruise speed.

[0159] For clarity, "wing-to-rotor skew" is measured relative to the wing orientation, which is fixed by the inherent attributes of the aircraft fuselage, regardless of the specific flight details. In this specification, with respect to an aircraft, directions and angles (and their changes) refer more specifically to directions / angles relative to the aircraft's pitch axis, unless otherwise specified. Yaw and roll components of direction / angles are explicitly indicated where necessary.

[0160] As used herein, the term "rotor" should be understood to include, from the outset, a unit that provides powered thrust to an aircraft. For example, it may include a propeller, a power source, and a shaft coupled to the propeller that the power source rotates, thereby rotating the propeller. The terms "motor" and "motor assembly" are also used herein in reference to elements capable of generating thrust. Generally, a motor (as a thrust-producing device) can be implemented as a rotor or another propulsion device, such as a turbine or jet. In this specification, the term "motor assembly" refers to a thrust-producing device in this sense. Conversely, when a rotor is referred to herein, it should be understood that the rotor is optionally replaced by another type of motor assembly (e.g., a jet or turbine), except when the description specifically relates to performance aspects specific to propellers, such as counter-rotating propellers and / or propeller pitch angle. It is understood that the orientation plane of a rotor's propeller is perpendicular to its net thrust direction. For embodiments using a propeller-less thrust generating device, references to the "plane of the propeller" can be interchangeably referred to as the "plane perpendicular to the net thrust direction."

[0161] When the term "motor" refers specifically to a power source such as an electric motor (and not the entire thrust-producing element / motor assembly), it will be specifically stated as "motor (per se)" or it will otherwise be indicated that the term refers specifically to the power source.

[0162] A power source may include, for example, an electric motor (per se), an internal combustion engine, or a turbine. Apparatus for supplying energy to a power source in the form of electricity or fuel is considered separate from the power source. Conversely, while reference to an aircraft's "propeller" implies the presence of a power source to which the propeller is attached via a rotating shaft, the term "propeller" as used herein does not include the power source per se. However, it does not exclude that a single power source may power two or more propellers, or conversely, that one propeller may be powered by two or more power sources. While the default convention of a one-to-one relationship between propellers and power sources is adopted for purposes of explanation herein, in some embodiments, for example, a single propeller is described as being operated by a compound power source, e.g., an electric motor transmitting power to a single propeller via two collinear stator / rotor pairs. In addition to the above meanings, "Rotor" is a term that is also used to refer to a component of an electric motor. This alternative meaning is indicated herein by, for example, specifying that the rotor is part of an electric motor and / or that the rotor is paired with a corresponding stator.

[0163] As used herein, the term "propeller orientation" refers to the direction of the plane in which the propeller rotates. Also, as used herein, the term "motor orientation" refers to the direction of the rotational axis of the propeller shaft (which is therefore perpendicular to the propeller orientation). A blade that is angled obliquely to the propeller orientation is therefore also angled to the motor orientation. However, for example, the blade may be angled 30° from the propeller orientation and 60° from the motor orientation. Specifically, the "oblique angle" of the blade relative to the propeller is optionally selected from the ranges of 5° to 45°, 10° to 35°, 15° to 30°, and 20° to 25°. The effects, including potential advantages, of selecting different angles are discussed in connection with the embodiments described herein.

[0164] Also, the components of the angles referred to herein are about the pitch axis unless otherwise noted.

[0165] Some of the specific embodiments described herein are described as having fixed wing-to-rotor skew angles. Herein, a description of an angle between directions (e.g., propeller and wing) as “fixed” functionally specifies that the angle does not change during operation of an aircraft embodiment, and specifically, does not change as part of the aircraft’s transition between forward flight and takeoff / landing modes. In some embodiments, the fixed angle between two components is provided structurally by the interconnection of the aircraft components (e.g., using static struts and fasteners) and does not provide for changes in the relative orientation of the components during flight. A fixed angle does not exclude being convertible by reconfiguring the aircraft on the ground. Examples with angles described as “fixed” do not exclude embodiments that allow members, such as struts, to flex to some degree during flight, e.g., as a result of changing forces on the aircraft. The term also does not exclude embodiments that include movable wing control surfaces. It is the body of the wing that is “fixed,” for example, because its orientation is set by the orientation at which the wing base is fixed to the aircraft fuselage. In some embodiments, "fixed" means immobile for at least 50% of the mass of the aircraft.

[0166] The angles discussed here in relation to the blade-to-rotor skew angle (which is the relative angle of the blade to the propeller plane) are also applicable to the blade-to-thrust axis skew angle (which refers to the rotor's thrust axis oriented perpendicular to the propeller). This angle is defined as 90° minus the fixed blade-to-rotor skew angle. When converted in this way, the angle still refers to the same actual configuration. When rotor thrust is deflected by a baffle, the "rotor's" thrust axis should still be considered to be perpendicular to the orientation of the rotor propeller plane.

[0167] Advantages of fixed and skewed blade-rotor angles include mechanical simplicity and reduced risk, cost, and / or weight for aircraft having this feature. However, it should be understood that there are potential advantages to blade-rotor skew angles, whether fixed or non-fixed. Thus, blade-rotor angles and blade-thrust axis angles described with respect to some embodiments as "fixed" may instead (in some alternative embodiments) optionally be implemented with the same specific angles as a preferred, but non-fixed, configuration available for use in forward flight. That is, the angles are variable during flight. Such embodiments have "variable" blade-rotor skew angles wherever explicitly mentioned. The teaching that such complementary embodiments should also be understood from the explicitly given examples of "fixed" angles does not alter the meaning of "fixed." Rather, any embodiment described herein as including fixed blade-rotor skew angles may be implemented with variable skew angles during pre-cruise flight. This is complementary to other embodiments that use variable blade-to-rotor splay for forward flight.

[0168] The wing-rotor skewing angle (fixed or variable) has potential advantages in some embodiments over the use (with fixed wings) of a fixed-pitch propeller design jointly optimized to provide thrust in both forward flight and hover modes of aircraft operation. Optionally, the rotor (or set of rotors) is operable in a first orientation (or set of orientations) relative to the wings that provides hover (e.g., net vertical) thrust and at least a second orientation (or set of orientations) relative to the wings that provides a forward flight thrust component. The orientation change (from hover mode to forward flight mode) can optionally be through any suitable angle. In some embodiments, the orientation change is through an angle substantially the same as the magnitude of the wing-rotor skewing angle described herein for a fixed wing-rotor skewing angle, e.g., an angle selected from at least one of the ranges of 5° to 45°, 10° to 35°, 15° to 30°, and 20° to 25°.

[0169] Optionally, the translation range is relatively small (up to about 5°, 10°, or other angle range). Optionally, this provides a potential advantage by ensuring that the rotor remains positioned in a safe and fully flyable orientation throughout its entire physically possible range of movement, while still preserving the potential benefits of additional hovering and / or flight efficiency gained by "tuning" the angle, for example, according to the aircraft's current loading and / or other flight conditions.

[0170] The angle conversion mechanism optionally includes, for example, any suitable gears, cables, and / or linked drive mechanisms. For certain types of angle conversion mechanisms, the use of small angle mechanisms may be advantageous. For example, a sliding bar-type linkage providing only a few degrees of actuation (e.g., less than 10°) may be lighter, simpler, less obtrusive, and more robust than a linkage allowing a full 90° of motion or even a smaller range of motion, such as 45°. Specifically, the rotor may be mounted using multiple link bars, each of which may be selectively tilted by shortening or lengthening relative to one another. A two-bar mount (for example) may achieve the same stiffness as a one-bar mount, but with less material. The power cabling and / or hose running to the rotor optionally provides the flexibility and slack needed for embodiments with variable-angle rotors.

[0171] Thus, in accordance with some of the above elements, in some embodiments, an unmanned or manned air vehicle is provided having three or more propellers and one or more wings, where the wings are parallel to the ground during flight and the motors are fixed at an angle relative to the wings so that in forward flight, some of the motor thrust generates lift and some pulls the air vehicle forward. During at least one of takeoff and landing, the air vehicle optionally hovers with the propellers parallel to the ground (i.e., the air vehicle is capable of vertical takeoff and landing).

[0172] Except with respect to aircraft embodiments having features that explicitly require a particular mode of operation, aircraft embodiments described herein are optionally manned or unmanned, or both, and the aircraft are optionally operated from the aircraft or the ground, or flown in an autonomous mode, or both.

[0173] Except for aircraft embodiments having features that explicitly require a particular motor type, the motor of the aircraft embodiments described herein is optionally at least one of, for example, an electric motor, an internal combustion engine motor, and a turbine motor (per se).

[0174] Except for aircraft embodiments having features that explicitly require a particular type of thrust propulsion, air thrust propulsion for the aircraft described herein is optionally provided by, for example, at least one of a propeller, a ducted fan, a jet engine, and a rocket booster.

[0175] An aspect of some embodiments of the present disclosure relates to an aircraft designed for both hovering and forward flight, in which at least 50% of the forward flight horizontal thrust is provided by rotors oriented at an angle to the wings. In some embodiments, the oblique angle is between about 20° and 25°. Optionally, the oblique angle is fixed. Optionally, the oblique angle is adjustable. Optionally, the oblique angle ranges between about 20° and 25°, with the oblique angle adjusted to another angle outside this range for forward flight.

[0176] The inventors have realized that although a canted rotor directs most of its thrust vertically (reducing its contribution to forward flight thrust), it may potentially regain otherwise lost horizontal thrust efficiency at high speeds, particularly for rotors with propeller blades that are also capable of generating efficient hovering thrust. Fixed propeller blades generally have an associated free-stream air speed at which they are most efficient. As the free-stream air speed increases beyond this speed, the blades lose their "mesh," and it is sometimes said that the blades cannot transfer much energy to the air passing through them because the air is already moving too fast. Higher blade pitch potentially increases high-speed performance. For example, variable-pitch blades are typically designed to increase their blade pitch at higher airspeeds.

[0177] Propeller blades, particularly for hovering applications, optionally have regions of relatively low blade pitch (to reduce / avoid stall at low freestream airspeeds), but this low pitch potentially reduces available power at relatively high freestream airspeeds (e.g., in cruising level flight). The more horizontally a rotor (e.g., a rotor with such blades) is oriented, the slower the propeller "sees" the airflow, since only a portion of the total forward airspeed is directed parallel to the rotor's thrust axis. This potentially allows a relatively hover-optimized (low-speed-optimized) propeller blade to nevertheless provide an efficient level of forward thrust at high airspeeds, e.g., about twice as fast (for a given efficiency) as an equivalent propeller oriented to generate fully horizontal thrust.

[0178] Without being bound by any particular theory of operation, the inventors have discovered that orienting a hover-capable propeller at an oblique angle of between about 20° and 25° relative to the forward flight direction potentially takes particular advantage of this propeller characteristic.

[0179] Optionally, a particular aircraft is designed for a particular optimum cruise speed. This is optimized, for example, by selecting at least one of wing / rotor angle, propeller blade pitch, and wing design within at least one constraint, such as weight, required range, required airtime, available energy budget, and required energy reserve. Typical cruise airspeeds for some embodiments of the present disclosure range from 30 to 130 kiloknots (approximately 55 km / h to 240 km / h). A typical magnitude of wing lift generated at approximately 50% of the design cruise speed is approximately 25% of the aircraft weight. A typical magnitude of wing lift generated at approximately 75% of the design cruise speed is approximately 50% of the aircraft weight. At cruise speed, a typical magnitude of wing lift generated is approximately 70% to 90% of the aircraft weight.

[0180] Before describing at least one embodiment of the present disclosure in detail, it is to be understood that the present disclosure, in its application, may refer, if necessary, to details of construction and arrangement of components and methods set forth in the following description and illustrated in the drawings. It should be understood that this is not necessarily intended to be limiting. Features described in this disclosure, including features of the present invention, are capable of other embodiments or of being practiced or carried out in various ways.

[0181] (Fixed wing-fixed motor mounting arrangement) Reference is first made to FIG. 1A, which is a schematic diagram of an aircraft with a rotor-driven wing 120 having a rotor 102 oriented at a fixed skew angle relative to the pitch axis direction 131 of the wing 120, in accordance with some embodiments of the present disclosure. Reference is also made to FIG. 1B, which is a schematic diagram of a propeller-driven wing 121 attached to an aircraft fuselage 140, having a rotor 102 oriented at a fixed skew angle relative to the direction 131 of the wing 121, in accordance with some embodiments of the present disclosure. These figures show the aircraft in horizontal forward flight, oriented in the direction shown. In both figures, a fixed wing-to-rotor skew angle 130 (about the pitch axis) between the propeller direction 132 and the wing direction 131 is shown. The angle shown is by way of example and, in some embodiments, is selected from at least one of the following ranges: 5° to 45°; 10° to 35°; 15° to 30°; and 20° to 25°.

[0182] 1A shows a pair of forward and aft rotors 102 secured to wings 120, with each rotor 102 in a pair secured by a respective forward arm 110A or aft arm 110B attached to the wings 120, such that the wings are located between the rotors 102. In FIG. 1A, as with the embodiments shown in other figures herein, in some embodiments each rotor 102 includes one or more propellers 104 and one or more motors 103 as a power source. In some embodiments, the rotors 102 include counter-rotating propellers, which is a potential advantage for balancing torque forces. Optionally, a fuselage (not shown) is attached to the wings 120.

[0183] 1B shows an aircraft configuration including a fuselage 140 to which at least one aft rotor 102 is attached by aft arms 110B at a fixed oblique angle relative to the wings 120, and a pair of forward rotors 102 are attached by respective forward / side arms 110H. Optionally, each rotor is equipped with multiple propellers and power sources.

[0184] 1A-1B includes a forward-mounted rotor 102 positioned below the wings 120 and an aft-mounted rotor 102 positioned above the wings 120. This configuration reduces or prevents airflow loss over the wings 120 (forward and aft) and into the rotors 102, even if all of the rotors were aligned along a forward-to-aft axis (parallel to the roll axis) (as shown in FIG. 1A). In full forward flight, both the rotors 102 and the wings 120 provide lift. To maintain altitude in hover and / or low-speed flight, power to the rotors 102 can be increased and / or the rotors 102 can be reoriented (e.g., by reorienting the aircraft, including reorienting the wings 120) so that the propeller heading 132 is, for example, substantially parallel to the ground.

[0185] Thus, the lift provided by wings 120 is not necessarily what keeps the aircraft airborne, but rather can provide extra lift (at the expense of wing drag) to offset the loss of lift (for any particular force) due to the forward tilt of rotor 102 thrust angle. Optionally, rotor 102-only operation is established as airworthy, so that the wings are not relied upon to support the entire weight of the aircraft. Optionally, this may allow wings 120 to be designed as smaller, thinner, and lighter structures, allowing for reduced drag and / or load.

[0186] In some embodiments, the rotors (including the motors themselves and the propellers) mounted forward of the center of gravity have different tilt angles (with respect to any axis, optionally including the pitch axis) and / or different propeller geometries relative to the rotors (including the motors themselves and the propellers) mounted aft of the center of gravity, this option being available in all aircraft embodiments described herein.

[0187] In some embodiments, an air vehicle (aircraft) is designed to place the center of lift generated by the wings forward or aft of the aircraft's center of gravity. This potentially introduces an upward or downward pitch moment to the air vehicle during forward flight. In some embodiments, this is compensated for by commanding the flight controls to, for example, increase power to the forward motors and decrease power to the aft motors if the center of lift is aft of the center of gravity, and vice versa if the center of lift is forward of the center of gravity. This option is available for all aircraft embodiments described herein, except for aircraft embodiments whose functionality requires the center of lift to be located at the aircraft's center of gravity.

[0188] Optionally, propeller and motor designs and / or tilt angles that are forward or aft of the center of gravity and require additional power during forward flight are optimized (e.g., by power source strength and / or propeller design) relative to forward flight performance, while propeller and motor designs and / or tilt angles that are forward or aft of the center of gravity and require less power during forward flight can be optimized for hover performance, potentially improving overall performance.

[0189] In some embodiments, optimizing the propeller design includes adjusting the (optionally fixed) pitch angle of the propeller blades to design a propeller that is relatively optimized for efficiency at higher relative velocities of passing air or at lower relative velocities of passing air. For example, the propeller may have maximum efficiency for a first (vertical) airspeed, but is intended for use in higher-speed flight. Some of the propeller's efficiency is regained in some embodiments by tilting the propeller at an oblique angle relative to the direction of forward flight. Thus, thrust efficiency lost in forward flight due to this oblique tilt is, optionally, at least partially regained.

[0190] In some embodiments, the pitch angle of a propeller blade optimized for forward flight performance has a pitch section that is so steep (i.e., produces such a large advance ratio for a given freestream airspeed) that it stalls during hovering flight. For example, the stall zone of a propeller blade optimized for forward flight is at least 20%, at least 30%, or at least 50% of the blade length. Typically, the stall zone extends from the radially inner portion of the blade (e.g., the side connected to the propeller hub). At a typical stall angle, the angle of attack deviates from the blade or chord direction by about 15° or more.

[0191] In some embodiments, for example, a propeller forward or aft of the center of gravity is provided with a relatively greater blade pitch tilt (and / or a larger stall zone) than a propeller on the opposite side. In some embodiments, a propeller located closer to or farther from the center of gravity is provided with a relatively greater blade pitch tilt (and / or a larger stall zone) than a propeller further from the center of gravity in pitch angle (propellers in one of these groups are optionally located on the same or different sides of the center of gravity in pitch angle). More specifically, in some embodiments, a propeller closer to the center of gravity is relatively optimized (e.g., by selection of pitch angle) to provide efficient high thrust during hovering compared to a propeller relatively farther from the center of gravity. The farther propeller optionally utilizes its relative mechanical advantage (e.g., stall losses, and / or by controlling the rotational speed of the propellers to provide low hover thrust.

[0192] In some embodiments, the forward and aft rotors are mounted at different distances from the center of gravity on the roll axis (e.g., different distances in hover flight and / or other flight angles). This potentially affects the relative thrust used to maintain a particular pitch angle. For example, torque exerted on the aircraft body by thrust from a rotor closer to the center of gravity is optionally balanced by torque exerted from a point further from the center of gravity by another rotor using lower thrust. In some embodiments, this is used to allow a rotor optimized for forward thrust to be angled (e.g., during landing) to generate less downward thrust than a rotor optimized to generate hover thrust. In some embodiments, by designing an aircraft with a center of gravity closer to the rotors that are more horizontal, the total forward thrust provided by rotors with more vertical in-flight thrust directions is reduced compared to the total thrust provided by rotors with more horizontal in-flight thrust directions. Alternatively, in some embodiments, the loss of vertical thrust from rotors with more horizontal thrust directions is compensated for by placing them relatively farther from the aircraft's center of gravity. Optionally, the flight control software is configured to adjust power output in response to differences in center of gravity, which may be caused by, for example, differences in load weight, differences in its distribution, passenger movement during flight, and / or cargo movement, and these adjustments are optionally explicitly programmed, or determined using, for example, automatic sensing of the responsiveness of the aircraft's attitude in response to commanded rotor power output.

[0193] It should be noted that in embodiments where the aerodynamic center of lift is offset from the aircraft's static center of gravity, the aerodynamic surface's center of lift may contribute to shifting the aircraft's "effective" center of gravity during forward flight.

[0194] In some embodiments, a relatively low-thrust "hover assist" motor is located relatively far from the center of gravity on the roll axis, e.g., at least two, three, or four times farther than its counterweight motor on the opposite side of the center of gravity. This offers the possibility of mounting the main motor on the same side of the center of gravity as the hover assist motor in a more horizontal thrust orientation, potentially benefiting forward flight efficiency. Optionally, the hover assist motor is recessed, carried within a streamlined cowling, featherable (e.g., flattenable in the direction of airflow), foldable, or otherwise retractable to reduce drag during forward flight. Optionally, the low power requirements make the hover assist motor and its mounting sufficiently lightweight that any weight and / or drag penalty is justified by the improved horizontal thrust efficiency. The hover assist motor, in some embodiments, does not need to be operational for the aircraft to be airworthy. Rather, it provides an optional alternative to slowing down, hovering, etc., which may be accomplished in some fashion when the aircraft is oriented at a different and / or more variable pitch angle.

[0195] It is particularly noted that if the forward and aft rotors are mounted with different pitch orientations, then in hover mode, neither propeller type will be parallel to the ground, but rather the propeller faces may be facing each other, for example, to form a V or inverted V. Such a configuration should now be considered to produce a thrust direction equivalent to "ground level" for rotors that all share a common pitch orientation, as far as net thrust direction is concerned, since the combined thrust generated by all rotors is perpendicular to the ground.

[0196] Reference is now made to FIG. 1C, which schematically illustrates airflows 152, 151, 150 passing through the forward rotor 102A, the aft rotor 102B, and over the wing 120, respectively, in an aircraft having rotors 102A, 102B oriented at a fixed oblique angle relative to the wing 120, in accordance with some embodiments of the present disclosure.

[0197] In some embodiments of the present disclosure, the motors (e.g., rotors 102A, 102B) are positioned such that the airflow to and from the propellers is not impeded by the wings or such that flow disturbances by the wings are minimized. For example, none of the motors direct air onto and / or toward the wing surfaces, none of the motors have thrust vectors toward the wing surfaces, or both. Furthermore, in some embodiments, the rotors are positioned such that they do not ingest air that may be disturbed by, for example, wing vortices. For example, as shown in FIG. 1C , the forward rotor 102A is attached below and forward of the wing 120 by arms 161 (it should be understood that at least one rotor is attached to each side of the aircraft), and the aft rotor 102B (also on each side of the aircraft) is attached above and aft of the wing 120 by arms 160. In addition to being located forward of the wing, the forward rotor 102A is located forward of the aircraft's center of gravity. In addition to being located aft of the wing, the aft rotor 102B is located aft of the aircraft's center of gravity.

[0198] During forward flight, the airflow over the wings 120 is not impeded (or only minimally impeded) by the wake of the propeller 104, and the airflow to and from the propeller 104 is not impeded by the wings 120. During hovering, the airflow to and from the propeller 104 is not impeded (or only minimally impeded) by the wings 120.

[0199] Avoiding this interference (e.g., mounting the rotors on arms that extend away from at least one of the planes of the fuselage 140 and the wing 120) is a potential advantage in terms of the efficiency of the lift and / or forward thrust generating components used in fixed-wing-rotor hybrid aircraft.

[0200] Reference is now made to FIG. 1D, which schematically illustrates the arrangement of landing gear 163, 164 for an aircraft having wings 120 and rotors 102 oriented at a fixed oblique angle relative to the wing direction 131, in accordance with some embodiments of the present disclosure.

[0201] In some embodiments, landing gear 163, 164 is attached to arms 161, 160 that support the motors (e.g., rotors 102A, 102B). The landing gear is optionally of the skid, leg, or wheeled type, for example, and is optionally retractable during flight.

[0202] Note that in the landing configuration shown, the angle 134 between the ground 165 and the pitch 132 of the propellers of rotors 102A, 102B is an oblique angle, and the angle 133 between the ground 165 and the pitch 133 of wings 120 is an oblique angle. In some embodiments, this will be the pitch direction of the landed aircraft, which is a direction between the hovering flight pitch of the aircraft and the full forward flight direction of the aircraft. The relationships, features, and potential advantages of different aircraft direction angles and / or relative orientations of individual elements are discussed herein, for example, in connection with Figures 10A-10F.

[0203] Reference is now made to Figures 1E-1G, which are schematic illustrations of a rotor-driven aircraft having a wing 120 and a rotor 102 oriented at a fixed oblique angle relative to the direction 131 of the wing 120, in accordance with some embodiments of the present disclosure.

[0204] 1E-1F each show a configuration in which pairs of forward and aft rotors 102 are secured to a wing 120, with each rotor 102 secured by a respective one of a forearm 110A or aft arm 110B attached to the wing 120, such that the wing is positioned between the pair of rotors 102. In some embodiments, each rotor includes one or more propellers 104 and one or more motors 103 as a power source. Any number of such pairs can be used at any location across the wing or its edge. In FIG. 1E, two pairs of rotors are attached to each of the tips (outer ends) of the wing 120, while in FIG. 1F, an additional pair of rotors is attached to the center of the wing 120. The difference between FIGS. 1F and 1E is the addition of two additional rotors. It should be understood that more rotors are optionally added (e.g., to support a heavier wing body), such as those shown.

[0205] FIG. 1G shows a pair of rotors 102 attached midway between the wings by bars 110A, 110B, in combination with an additional pair of rotors 102, each attached to each outer tip of the wing 120 by a bar 110C, with the bars 110C extending approximately parallel to the pitch axis of the wing 120, rather than forward or aft.

[0206] Reference is now made to Figures 1H-1O, which are schematic illustrations of a rotor-driven aircraft having wings 121 and a fuselage 140, with rotors 102 oriented at a fixed oblique angle relative to the direction of the wings 121, in accordance with some embodiments of the present disclosure.

[0207] The motors are optionally mounted directly on any suitable portion of the aerial vehicle, such as the fuselage or fuselage extensions (including structures also referred to herein as "arms" or "bars"), wings, struts between wings or between wings and fuselage, and vertical surfaces of stabilizers.

[0208] In some respects, the overhead swept-wing design of Figure 1H is similar to the design shown in Figures 11A-11C, with two rotors 102 attached to arms 110A partially below and forward of fuselage 140 (and completely below wings 121) and two rotors 102 attached to arms 110B above and aft of fuselage 140. Wings 121 are attached overhead to fuselage 140. The forward arms may optionally be two separate arms or may be one single bar, such as a bar across the nose.

[0209] 1I-1M show an example configuration in which several motors 103 and propellers 104 are attached to a vertical stabilizer 110C.

[0210] 1N-1O show examples of configurations where motors are attached to bars 110E, 110G between wings 121, 124, and optionally to extensions of these bars in a multi-wing or canard configuration. Arms connect the two wings or a wing and a stabilizer, and motors can be attached to the arms between the two surfaces.

[0211] Reference is now made to FIG. 1P, which is a schematic diagram of a rotor-driven aircraft including a wing 121 and two fuselages 140A, 140B, with rotors 102, 102C oriented at a fixed oblique angle relative to the orientation of the wing 121, in accordance with some embodiments of the present disclosure. The aircraft of FIG. 1P includes seats 150 for two passengers in each of the two fuselage sections 140A, 140B. Note that there is no specific requirement that the rotor propellers be the same size. For example, propeller 104C of rotor 102C is optionally larger in size than propeller 104 of rotor 102.

[0212] Reference is now made to FIG. 1Q, which illustrates a wing 121 and a four-seat fuselage in accordance with some embodiments of the present disclosure. 1 is a schematic diagram of a rotor-driven aircraft with fuselage 140C, 140B, with rotors 102 oriented at a fixed oblique angle relative to the direction of wings 121. Details of rotors 102 and their mounting struts have been omitted to show the arrangement of four seats 150 within fuselage 140C.

[0213] 1E-1Q include embodiments with up to six rotors. It should be understood that more rotors are optionally added as dictated by the size and / or weight of the aircraft (e.g., to support a heavier aircraft). For example, in some embodiments, the fuselage 140 of FIG. 1O is elongated, and additional rotors 102 are added along bar 110G as needed. The additional rotors 102 do not have to be added along the same bar; for example, they could be added to additional struts or bars. It should also be emphasized that not all rotors 102 are required to be coplanar or share the same diagonal orientation relative to the wings 121, 124. For example, different orientation configurations are described in connection with FIGS. 10A-10I.

[0214] It should be understood that the mounting options shown are applicable to airborne vehicles having tails, canards, multiple wings, or other stabilizing or lift-generating elements, including configurations of elements not explicitly shown.

[0215] (Relative orientation and attitude changes of elements during maneuver) Reference is now made to Figures 10A-10C, which are schematic illustrations of aircraft attitude changes during takeoff and / or landing operations, in accordance with some embodiments of the present disclosure.

[0216] 10A depicts an aircraft 1015 on the ground. In this example, both the forward rotor 1022A and the aft rotor 1022B are oriented at the same oblique pitch angle relative to the main wing 1025 and optional secondary wing 1026 (optionally implemented as part of the aircraft's tail). Optionally, they are oriented at different pitch angles, but the same pitch angle is used here for purposes of illustration.

[0217] Also shown are aft and forward landing gear carriages 1023, 1024, fuselage 1025, and variable-position ("reclining") cockpit chairs 1021. Cockpit chairs 1021 generally represent an arrangement within an aircraft cabin that allows for tilt adjustment of passengers and / or cargo independent of the overall pitch of the aircraft.

[0218] A schematic chair 1001 corresponds to the position of the cockpit chair 1021. A wing surface 1002 corresponds to the pitch direction of the main wing 1025 (optionally parallel to the pitch direction of the secondary wing 1026). A propeller surface 1003 represents a plane parallel to the propeller direction of the rotors 1022A, 1022B. A ground contact surface 1004 represents a surface on which the landing carriages 1023, 1024 touch the ground. The landing gear carriages 1024, 1023 are optionally movable between multiple positions (e.g., retracted during flight), but are shown fixed for illustrative purposes.

[0219] In the example shown, the wing-to-rotor pitch angle is shown at 45°. When the aircraft 1015 is landing, the landing gear carriages 1023, 1024 are configured to hold the aircraft so that the wing planes 1002 are at an angle to the ground (approximately 15°) and the propeller planes 1003 are at an opposite angle to the ground (approximately 30°). The schematic chair 1001 is shown in a vertical position. Optionally, the schematic chair 1001 is not strictly vertical, but rather a "comfortable" vertical, for example, in terms of suitable visibility and / or ergonomics provided to cabin occupants.

[0220] FIG. 10B shows that all the same elements and the same relative angles are used, and the entire aircraft is pointing upward (nose up) during at least one of the periods of hovering, initial takeoff, and landing contact. The cockpit chair 1001 is shown pitched approximately 30° from the horizontal position during takeoff. The propeller surfaces 1003 are oriented horizontally, thereby directing the thrust vectors 1010 (corresponding to thrust from rotors 1022A, 1022B) vertically downward (corresponding to the generation of upward lift). The wing surfaces 1002 are tilted away from the effective direction for providing lift during flight and are not critical to flight characteristics during hover. The schematic chair 1001 is tilted backward by approximately 30°. For comparison, a typical airliner's climb angle is optionally approximately 15-20° (though some commercial aircraft typically climb at 25°). Optionally, the cockpit chair 1021 can tilt itself forward from the position shown during takeoff, for example, up to approximately 15°, remaining fully within the range of a typical airliner's climb angle.

[0221] 10C shows the same elements again, with most of the relative angles remaining the same, except now the schematic chair 1001A (and corresponding cockpit chair 1021) is shown tilted forward approximately 15° from its previous orientation (again represented by the dotted line showing the schematic chair 1001). The aircraft 1015 is oriented for full forward flight, tilted downward approximately 30° from the landing attitude. The rotors 1022A, 1022B are oriented at an angle of approximately 45° to the ground, with their thrust split approximately evenly between horizontal thrust 1012 and vertical thrust 1011. The forward air velocity of the wings 1025 provides lift (opposite the downward push vector 1013) and a small amount of drag (opposite the forward push vector 1014).

[0222] In particular, it can be seen that cabin occupants optionally experience pitch changes ranging from +15° to −30° (without chair or other cabin adjustments), and only from approximately 0° to −30° with adjustments (as shown). Chairs with adjustments greater than ±30° can further reduce this range, for example, to between 0° and −15°. Furthermore, even though rotors 1022A, 1022B may pitch up to a total of approximately 45° during operation, typical landing and cruise forward flight attitudes, in particular, are optionally maintained at 0°, with periods of maximum pitch limited to takeoff and / or landing. Furthermore, the transition from landing to forward flight optionally occurs at an even more limited angle, i.e., a direct change from the landing orientation of FIG. 10A to the forward flight orientation of FIG. 10C, which is tilted approximately 15° further forward (nose down). This may result in passengers not experiencing any attitude changes (e.g., if the chair tilt is synchronized with the forward flight pitch transition). This is particularly true, but not exclusively, in aircraft with only a relatively small number of passenger seats (e.g., one, two, or three rows), where there are height differences between seat rows in the aircraft's landing configuration, e.g., differences of about 20-50 cm or more. In longer aircraft (e.g., four or more rows of seats), this may be inconvenient, and a more level landing configuration of seats is optionally used. In hover flight, the seating area of ​​such long aircraft may be reoriented (along with the aircraft itself), resulting in, for example, the aisles between seats in the cabin floor being tilted backward. In forward flight, the same cabin floor is tilted forward.

[0223] Optionally, forward attitude changes experienced by the cabin occupants / cargo during pre-landing deceleration are kept to a minimum by at least one of slow deceleration, utilizing aerodynamic drag, slowly reducing forward pitch, and yawing the entire aircraft backward during part of the deceleration phase. A descent to landing from a full hover would typically still subject the occupants / cargo to a pitch of approximately -30° (in this example). In some embodiments (e.g., using rotary landing gear), a slow ground speed (optionally astern) descent for landing may still reduce this experienced pitch in situations where such a landing method is deemed appropriate.

[0224] Efficiency can potentially be increased if the wing lift (opposite vector 1013) dominates the downward rotor thrust vector 1011, so that more motor thrust is directed toward the horizontal thrust vector 10 Note that the rotors 1022A, 1022B are oriented toward the wing 12. Thus, potential benefits are achieved by accepting extra crew lean during any phase of flight in exchange for a more forward-leaning flight orientation of the rotors 1022A, 1022B.

[0225] 12A, which is a flow chart that schematically illustrates a method for reorienting an aircraft during takeoff of the aircraft, according to some embodiments of the present disclosure. The illustrated aircraft (e.g., as described with respect to FIGS. 1A-1B, 10A-10F, and / or 11A-11C) includes a wing and a rotor oriented at a fixed skew angle relative to the pitch direction of the wing 120.

[0226] In block 1210, in some embodiments, the aircraft is loaded / boarded while the aircraft is on the ground and the aircraft's propellers are oriented obliquely to the ground, as described in connection with, for example, FIG. 1C and FIG. 10A. The angle of the propellers is optionally selected, for example, within a range of approximately 10° to 35°, 15° to 30°, or some other angle range. Optionally, the aircraft's wings are also oriented obliquely to the ground, for example, within a range of approximately 10° to 35°, 15° to 30°, or some other angle range. Optionally, the propellers and wings are oriented at opposite oblique angles to the ground, such that the sum of their respective oblique angles to the ground provides an oblique angle relative to one another.

[0227] At block 1212, in some embodiments, the first stage of takeoff occurs. This stage optionally includes reorienting the aircraft while a portion of the landing gear remains both in contact with and stationary relative to the ground. In some embodiments, this includes tilting the aircraft backward until the aircraft's propellers, initially oriented at an angle to the ground, are oriented substantially parallel to the ground. In some embodiments, this also results in an increase in the pitch angle of the wings relative to the ground.

[0228] At block 1214, in some embodiments, the second phase of takeoff occurs. The aircraft lifts off the ground and hovers. If the aircraft was not reoriented while on the ground as described in block 1212, it optionally reoriented in the air after a brief period of lateral acceleration. During hovering flight, the aircraft optionally climbs to any suitable altitude for transitioning to forward flight. Block 1214 is also optional for some embodiments of the invention; in some such embodiments, takeoff proceeds directly from the landing state to forward flight, avoiding hovering. However, a potential advantage of gaining altitude during hovering is at least one of reducing the risk of collision with nearby ground objects and providing an opportunity to reoriented the aircraft before commencing forward flight, which could potentially avoid the need for a large turning maneuver after forward speed is achieved.

[0229] At block 1216, in some embodiments, forward flight follows. This transition to forward flight, in some embodiments, includes adjusting the power ratio between the aft rotor and the forward rotor (or other thrust device), causing the aircraft to bank and accompanying reorientation of the net thrust component toward the horizontal. The transition to forward flight optionally occurs when the aircraft is constantly supported by downward rotor thrust, at a power regime sufficient to keep the aircraft lifted (e.g., prevent downward acceleration, maintain altitude, or both). As forward speed increases, the wings generate lift. Optionally, the ratio of forward to aft rotor power is adjusted to maintain a desired pitch. Optionally, upon reaching a desired altitude, for example for cruise forward flight, rotor power is reduced until airflow over the wings provides at least a portion of the lift required to maintain altitude. This potentially improves efficiency over using rotor thrust alone to provide vertical thrust.

[0230] 12B, which is a flow chart that schematically illustrates a method for reorienting an aircraft upon landing of the aircraft, according to some embodiments of the present disclosure. The illustrated aircraft (e.g., as described with respect to FIGS. 1A-1B, 10A-10F, and / or 11A-11C) includes a wing and a rotor oriented at a fixed oblique angle relative to the pitch direction of the wing 120.

[0231] In block 1220, in some embodiments, the aircraft is in forward flight, for example, as described in block 1216 of FIG. 12A.

[0232] At block 1222, in some embodiments, forward speed is canceled during the first phase of landing. This is done, optionally using one or more of several techniques, for example, as described in connection with FIGS. 10A-10C. Often, canceling forward speed involves orienting the rotor pitch to reduce or redirect the horizontal thrust component. In some techniques, the rotor may additionally or alternatively be reoriented by yawing the entire aircraft. Canceling forward speed is optional in some embodiments, such as when performing a rolling landing.

[0233] In block 1224, in some embodiments, during the second phase of landing, the aircraft cancels its horizontal velocity and descends to a hover, with the propeller direction generally parallel to the ground. If the propellers are pointed in a different direction (e.g., by pitching), their net horizontal thrust vector is controlled to zero, allowing the aircraft to remain stationary.

[0234] In block 1226, in some embodiments, the aircraft resumes the state described in connection with block 1210 of Figure 12A.

[0235] Reference is now made to Figures 10D-10F, which are schematic illustrations of an aircraft 1035 having a forward rotor 1032A and an aft rotor 1032B at different angles relative to one another, in accordance with some embodiments of the present disclosure. Reference is also made to Figures 10G-10I, which are schematic illustrations of an aircraft 1055 having a forward rotor 1052A and an aft rotor 1052B at different angles relative to one another, in accordance with some embodiments of the present disclosure. In Figures 10D-10F, the rotors are angled relative to one another so that their planes intersect relatively below or below the aircraft. In Figures 10G-10I, the rotors are angled relative to one another so that their planes intersect relatively above or above the aircraft. Other elements in Figures 10D-10I are optionally identical to those described with respect to, for example, Figures 10A-10C. The relative angle is optionally selected within a range of up to about 7° (e.g., 5°). This tilt is sufficient to provide potentially significant performance tuning benefits (optionally differential tuning of the forward and aft motors for hover / forward flight performance) while maintaining sufficient similarity of tilt to facilitate coordinated operation of the motors in both hover and forward flight modes. Optionally, the relative tilt is selected within different ranges, for example, up to about 15°, 30°, or other angles, although greater differences within these ranges may result in increased control complexity and reduced performance payoff for one or more flight modes, depending on the performance characteristics for which the aircraft is optimized.

[0236] 1A-1B, the forward rotors 1032A, 1052A and aft rotors 1032B, 1052B can be oriented at different fixed angles (corresponding to pitch planes 1042, 1043, 1052, 1053, respectively) relative to each other, as well as relative to the wings 1025 and wing surfaces 1002. Optionally, the rotors can be oriented at least in a landing state (FIGS. 10D, 10G), a hovering state (FIGS. 10E, 10H), and full forward flight (FIGS. 10F, 10I). The transition angles between them are as small as those for the corresponding same orientation configurations as shown in Figures 10A-10C.

[0237] Optionally, both sets of rotors 1032A, 1032B, 1052A, 1052B operate at equal power and opposite pitch angles (relative to horizontal) during hovering flight, optionally making equal contributions to the hovering thrust so that their respective vertical thrusts cancel each other out (e.g., as shown in FIGS. 10E and 10H). This hovering condition potentially differs in control characteristics from a configuration in which all rotors are vertical, because adjustments to the power themselves, and attitude adjustments influenced by the power, affect horizontal thrust while attempting to hover. The control programming and / or electronics are preferably configured to take this into account. Optionally, different power and / or different opposite pitch angles are used during hovering. Optionally, the rotor sets on the same side of the center of gravity along the roll axis include rotor sets configured in different orientations (e.g., the two outer rotors in one orientation and the middle rotor in another orientation). The net direction of the normal vector is also controlled by differentially powering the rotors. To simplify the control inputs for such configurations, fly-by-wire control logic is optionally used.

[0238] 10F and 10I show aircraft 1035, 1055 in a fully forward flight attitude. Rotor 1032B (rear in FIG. 10F) or rotor 1052A (forward in FIG. 10I) is oriented to generate more horizontal thrust, while rotors 1032A, 1052B are oriented to generate more vertical thrust. Depending on the amount of lift provided by wing 1025, rotors 1032A, 1052B are optionally significantly throttled, e.g., reduced to power sufficient to control and / or maintain level flight. In contrast, nearly the entire thrust of rotor 1032B or 1052A is available to provide forward thrust. Note that maintaining a slight canted angle relative to the direction of forward flight offers the potential advantage of maintaining a lower forward thrust ratio and potentially greater efficiency at high speeds than if the rotors were oriented completely perpendicular to the direction of forward flight.

[0239] If the aircraft 1035, 1055 is accidentally rotated into a forward flight pitch (or due to a stall in the wings 1025) before sufficient lift is generated from the wings 1025 to keep the aircraft afloat, the aircraft 1035, 1055 can be returned to a stable flight mode, for example, by increasing power to the rotors 1032A, 1052B and optionally swinging the pitch angle back so that the rotors 1032B, 1052A can also contribute more thrust in the direction of lift.

[0240] Optionally, restoration of sufficient rotor lift (optionally consisting of lift from the rotor alone sufficient to at least one of maintain altitude and prevent downward acceleration) is accomplished from full forward flight pitch (with wings level to provide flight lift) without requiring a simultaneous angle change, or without requiring an angle change of more than about 5° in pitch angle, for example. This does not exclude further angle changes that could potentially occur as rotor thrust balances the aircraft in the new equilibrium pitch. The "without requiring" qualification refers to the level of the instantaneous vertical component of thrust being provided sufficient to balance the weight of the aircraft and at least one of loss of altitude and prevention of downward acceleration.

[0241] In some embodiments, rotor thrust is adjusted in a forward flight direction to mitigate wing stall without a change in orientation from at least one of a zero airspeed condition, a low airspeed condition (e.g., less than about 30 km / h), and an airspeed condition that is within 10-20 km / h below the wing stall speed. The rotor pairs produce a net thrust vector pitch canted relative to one another by an angle of less than 45°, less than 30°, or less than 15°.

[0242] (Manned Aircraft Design) Reference is now made to Figures 11A-11E, which are schematic illustrations of an aircraft 1100, optionally configured for manned flight, according to some embodiments of the present disclosure, having wings 1120 and rotors 11012 oriented at a fixed oblique angle relative to the pitch direction of the wings 1120.

[0243] In some embodiments, aircraft 1100 optionally includes upswept wing 1121 with winglets 1122 at its leading edge and optionally attached to the top of fuselage 1140. Passenger / crew compartment 1141 (e.g., a two-person compartment) optionally includes a transparent canopy with a large visibility angle and optionally low and flared near the nose of the aircraft to increase downward and / or forward visibility at aft pitch angles such as those expected during hovering. In some embodiments, the low and flared canopy is facilitated by concentrating instruments near the centerline, e.g., using panel display 1142 as the instrument display and / or user interface.

[0244] The aft rotor 1102 is attached to an arm 1110B that extends aft and above the wing 1121. The forward rotor is attached to an arm 1110A that extends forward and below the wing 1121, optionally with the center of the forward rotor 1102 (e.g., the center of the electric motor 1103) located below the fuselage 1140. A high-wing design has the potential advantage of allowing the lower rotor to be mounted not too far below the fuselage 1140. For example, the blades of the rotor propeller 1104 can be selected to extend above the bottom of the fuselage 1140.

[0245] In some embodiments, a landing gear carriage 1162 is attached to the forearm 1110A. The aft landing gear carriage 1161 is attached to the fuselage 1140, optionally by a dedicated strut.

[0246] Entrance to aircraft 1100 is optionally via a door defined around the aircraft's forward cockpit window 1143, which is optionally hinged, for example, near the front of the aircraft or along the centerline, and which flips up when opened. Additionally or alternatively, in some embodiments, entrance to aircraft 1100 and / or a cargo hold of aircraft 1100 is located behind and below fuselage 1140, for example along underside 1144. Optionally, underside 1144 is configured as door 1145 ( FIG. 11D ), which pivots downward to open and, optionally, forms a ramp and / or staircase. The treads of such staircase are oriented horizontally (parallel to a plane defined by the landing gear contact surfaces) when door 1145 is in its normal open configuration and aircraft 1100 is fully on the ground.

[0247] Additionally or alternatively, entrance to aircraft 1100 is provided in a configuration that includes a side door entrance, such as stairway 1147 (which optionally folds away from the fuselage), and / or doorway 1148 (illustrated as opening in a raised configuration, but doorway 1148 optionally opens in a different direction). The treads of stairway 1147 are oriented horizontally when aircraft 1100 is in its normal landing configuration. Note that stair access is optionally provided in any of the passenger-sized aircraft embodiments described herein. Typically, the stairway treads are configured such that when the aircraft is fully landed, the treads are deployed with the treads parallel to the ground. In some embodiments, this results in a configuration in which at least one of the treads and wings and / or the treads and rotors are oriented at an angle relative to the other. do.

[0248] The relative oblique pitch angle of the rotor 1102 and blades 1121 can be seen in any of Figures 11A-11C. In Figure 11C, a plane 1160 is shown oriented parallel to the plane of the propeller 1104 and obliquely intersecting the blades. In the embodiment shown, this plane also intersects each of the rotors 1102 at approximately the same location. This is an optional feature; in some embodiments, the rotor planes are offset therefrom and parallel. Optionally, the rotor planes are also oblique to one another, for example, as described with respect to Figures 10D-10I herein.

[0249] Reference is now made to FIG. 16, which is a schematic illustration of an aircraft 1600, optionally configured for manned flight, according to some embodiments of the present disclosure, having wings 1621 and rotors 1602 oriented at a fixed oblique angle relative to the pitch axis direction of the wings 1621.

[0250] In some embodiments, the auxiliary rotor 1605 is attached to the blade 1621, optionally at an oblique angle relative to the blade. This oblique angle may be the same as the rotor-to-blade oblique angle of the rotor 1602, or may be a different oblique angle, for example, an oblique angle adjusted to account for partial thrust vectors due to airflow from the rotor 1605 acting on the blade 1621.

[0251] In some embodiments, the auxiliary rotors 1605 operate to provide additional thrust as needed for hovering, horizontal maneuvering, acceleration, deceleration, and / or counter-thrust balancing (particularly during landing and / or takeoff, and optionally in the event of failure of one or more of the rotors 1602). In the configuration of Figure 16, the rotors 1605 are particularly suited for use in providing auxiliary hovering thrust, as using them in full forward flight would tend to apply thrust downward directly onto the plane of the wings 1621. Accordingly, the propellers of the auxiliary rotors 1605 are optionally configured with blade pitch angles suited to low airspeeds.

[0252] Optionally, auxiliary rotors 1605 are configured to at least one of feather, sweep back, and pivot their blades out of the airflow for drag reduction during high-speed forward flight. During full forward flight, the propeller blades of rotor 1605 are optionally held in a position that minimizes the aerodynamic impact of wings 1620. For example, the propeller blades of rotor 1605 are held in their most vertical position, perpendicular to the adjacent plane of wings 1621, or both. Optionally, two-bladed propellers are used on rotor 1605. While this may result in increased vibration from these rotors during high-speed forward flight, the impact of the vibrations on, for example, cabin comfort is optionally minimized by operating these rotors during takeoff, landing, and / or emergency.

[0253] Auxiliary rotor 1605 is optionally lighter than rotor 1602 (e.g., a motor of about half the power and weight, a propeller with fewer blades, e.g., two blades versus four), thereby reducing its weight penalty. Optionally, auxiliary rotor 1605 has relatively shorter propeller blades than rotor 1602 (potentially reducing parasitic drag in forward flight). Also, optionally, in operation, it can rotate at a higher revolutions per minute to help compensate for its smaller blade size.

[0254] Optionally, auxiliary rotors 1605 are operated to generate extra thrust during forward flight, and optionally used in reverse, for example, during landing maneuvers. In some embodiments, auxiliary engines are a convenience and / or safety factor, but are not necessary to operate the aircraft or otherwise provide normal airworthiness. In some embodiments, The auxiliary engines 1605 provide redundancy, for example, for stabilization. For example, an event such as a collision (e.g., a collision with a tall ground obstacle damaging at least one of the rotor attachment arms 1610A, 1610B and one of the rotor attachment arms carrying the rotors) could potentially disable the main motors 1602 and create flight instability that cannot be fully compensated for by the remaining rotors 1602. In such a condition, the auxiliary rotors 1605 can optionally be activated to help maintain at least some flight stability, for example, to balance the aircraft long enough to execute an emergency landing. Typically, a four-rotor configuration would be extremely difficult to stabilize in a hover if reduced to a three-rotor configuration. On the other hand, five, six, or more rotors generally provide easier stabilization in the event of a rotor loss.

[0255] With respect to energy balance considerations, the auxiliary rotors 1605 potentially provide an advantage by reducing the hover thrust generation requirements for at least some of the rotors 1602. These rotors can thus be designed to be more efficient in forward flight (e.g., with a larger blade pitch) and / or can operate at lower power during hover flight (e.g., a smaller propeller stall zone). Optionally, the propellers of the auxiliary rotors 1605 are rotated parallel to the wings and retracted when the aircraft lands. In a two-bladed propeller configuration, this potentially reduces interference of the auxiliary rotor blades with ground equipment and / or obstacles.

[0256] Although aircraft 1600 is shown without landing gear, landing gear is optionally provided, for example, as shown in Figures 11A-11C or in another configuration. Compared to the aircraft of Figures 11A-11C, aircraft 1600 has vertical stabilizing surfaces located on vertical stabilizer 1622 rather than wingtip stabilizers. In some embodiments, vertical stabilizer 1622 provides sufficient aerodynamic stability to aircraft 1600 in forward flight, so that positive active flight control for yaw stabilization (e.g., by adjusting rotor-propeller speed and / or rotor thrust) is optionally deactivated for forward flight.

[0257] Fuselage 1640 and passenger / crew compartment 1641 optionally correspond to fuselage 1140 and passenger / crew compartment 11140.

[0258] (balance tilt motor mounting subframe) Reference is now made to Figures 2-4, which schematically illustrate rotor orientation transition mechanisms for an aircraft 200, according to some embodiments of the present disclosure.

[0259] In some embodiments, wings 1B are pitched parallel to the ground during flight (wings-horizontal) and are fixed to fuselage 1A to form wing-fuselage assembly 1. One or more motors (e.g., rotors) 8 are fixed to a subframe 2, which is attached to wing-fuselage assembly 1 by hinges 3, allowing them to rotate about the pitch axis of the aircraft.

[0260] In some embodiments, the subframe 2 rotates freely at the hinge 3, but a friction device limits the speed of rotation.

[0261] In some embodiments, the rotation angle of the subframe 2 is mechanically limited to a maximum pitch down angle relative to the wing and fuselage 1, which aligns the thrust of the motors 8 with the forward axis 6 of the aircraft, although this angle may be less than or greater than the exact forward axis direction.

[0262] The rotation angle of the subframe 2 is, in some embodiments, mechanically limited to a maximum pitch-up angle relative to the wings and fuselage, where the thrust of the motors 8 is directed upward and vertically relative to the aircraft 200. It is aligned slightly forward of axis 5.

[0263] The longitudinal position of the hinges on the fuselage and wing is aft of the center of gravity 7 of the fuselage-wing assembly 1.

[0264] The longitudinal position of the hinge 3 on the subframe 2 is, in some embodiments, forward of the center of gravity of the subframe 2. In high speed forward flight, the center of lift 4 of the fuselage-wing assembly is forward of the center of gravity.

[0265] During slow flight, takeoff, landing, and hovering (FIG. 2), the lift force 13 from wing 1B is small and the balance between gravity 10 and the vertical lift force 9 of the motor on hinge 3 dominates. This causes the fuselage-wing assembly 1 to rotate in a pitch-down direction 11 until it is in equilibrium at its maximum pitch-down orientation relative to the subframe 2.

[0266] In this orientation the subframe 2 generates lift which causes the aircraft to hover with the longitudinal axes of the fuselage 1A and wings 1B at a slight pitched up angle 15 relative to the ground.

[0267] In hover, in some embodiments, the stability and control of the aircraft is controlled using differential thrust to the motors, which is coordinated by an electronic flight controller.

[0268] To begin forward flight, the entire aircraft 200 uses the actuation thrust control to pitch down to an angle where the longitudinal axes of the fuselage 1A and wings 1B are parallel to the ground. In this orientation, a portion of the thrust of the motors 8 is directed forward, and the aircraft 200 begins to fly forward.

[0269] As forward speed increases (Figure 3), the lift force 13 generated by the fuselage-wing assembly 1 becomes stronger, and the torque exerted by the center of lift on the hinge 3 becomes greater than the opposing torque exerted on the hinge 3 in static balance with gravity. At this point, the pitch-up torque exerted on the hinge 3 by the fuselage 1A and wing 1B causes a pitch-up rotation 12 of the fuselage 1A and wing 1B relative to the hinge.

[0270] In some embodiments, a hinge friction device optionally slows this rotation.

[0271] The flight controller is configured to sense hinge rotation, for example, by sensors on the hinge 3 and orientation sensors on the subframe 2 and / or fuselage-wing assembly 1. The flight controller optionally applies differential thrust 14 to change the orientation of the subframe 2 to compensate for and eliminate pitch up of the fuselage and wings.

[0272] Figure 4 shows the configuration at cruising speed where the balance of forces causes the subframe to pitch down fully to the mechanical stop 6.

[0273] The process of slowing down is the opposite of speeding up.

[0274] During forward flight, the attitude of the aircraft is optionally controlled by differential thrust and / or control surfaces.

[0275] In some embodiments, the flight control system includes sensors to detect the attitude (pitch orientation) of both the fuselage-wing assembly 1 and the subframe 2, and the angle between them. This can be done by sensors on the hinge 3, orientation sensors on the subframe 2 and / or the fuselage-wing assembly 1.

[0276] In some embodiments, the flight controls are configured to command differential thrust to maintain a target attitude (pitch direction) of either the fuselage-wing assembly 1 or the subframe 2, or an angle that is a function of a combination of both attitudes.

[0277] Optionally, the aircraft comprises an additional motor fixed to the fuselage-wing assembly. In some embodiments, the additional motor has an upward thrust perpendicular to the forward flight axis and is used only for hovering and vertical takeoff and landing. Optionally, the additional motor is positioned at an angle to the orientation of the wings 1B, for example, as described with respect to Figures 1A-1O.

[0278] Additional motors are optionally used to control the attitude of the fuselage-wing assembly. The hinges may be frictionless, especially if auxiliary motors are used. Conversely, one or more slew motors are optionally used primarily to adjust the attitude of the fuselage-wing assembly. This use is optionally activated at selected moments in flight, for example, during landing maneuvers, where the use of auxiliary thrust capability is an optional alternative to variable pitch maneuvers. Off-path turns (e.g., horizontally) during forward flight can result in drag savings. Turns are optionally performed by operating the slew motors in forward or reverse thrust mode to move the motors to their angular limits.

[0279] Motors (per se) for the aircraft 200 include, but are not limited to, electric, internal combustion, and / or turbines, and air thrust propulsion for the aircraft includes, but is not limited to, propellers, ducted fans, jet engines, and / or rocket boosters.

[0280] (Subframe tilt for independent thrust control) Reference is now made to Figures 5-6, which schematically illustrate a rotor orientation transition mechanism for an aircraft 500, according to some embodiments of the present disclosure.

[0281] In some embodiments, manned or unmanned aerial vehicle 500 includes four or more propellers and one or more wings, with wings 501B configured to pitch parallel to the ground to provide lift during flight and secured to fuselage 501A. Portions of motors 504 are secured to fuselage-wing assembly 501 so that the thrust of these motors is aligned generally upward along the vertical axis of the aircraft (although this angle is optionally less than or greater than exactly upward). Another portion of motor 505 is secured to subframe 502 attached to wing-fuselage assembly 501 by hinge 503, allowing subframe 502 to rotate about the pitch axis of the aircraft.

[0282] Subframe 502 is free to rotate on its hinges up to a maximum pitch-up angle that is mechanically limited relative to wings 501B and fuselage 501B, thereby generally aligning the thrust of motors 505 with the forward axis of aircraft 500. Optionally, this angle is less than or greater than the exact forward axis direction of the aircraft.

[0283] The rotation angle of subframe 502 is limited to a minimum pitch down angle that is mechanically limited relative to wings 501B and fuselage 501A, at which pitch angle the thrust of motors 505 is aligned approximately upward along the vertical axis of aircraft 500. Optionally, this angle is less than or greater than the exact upward direction.

[0284] The hinges 503 on the subframe 502 are positioned so that when the subframe motor thrust is upward, one or more motors 505 are forward of the hinge and one or more motors are aft of the hinge. The flight controller monitors the sensors on the hinges, the subframe and the fuselage-wing axis. The hinge rotation is sensed either by an orientation sensor on the subframe assembly, or a combination of these. The flight controller can apply differential thrust to the subframe motors forward and aft of the hinge to change the orientation of the subframe.

[0285] During takeoff (FIG. 6), landing, and hovering, in some embodiments, the flight controller applies differential thrust to the subframe motors to maintain a generally upward thrust direction for the subframe motors 505. This is done, for example, by applying more thrust to the motors 505 forward of the subframe hinges 503 and less thrust to the motors 505 aft of the subframe hinges 503, or by applying any relative thrust that "pins" the subframe 502 at its minimum pitch down angle (and / or closest to horizontal).

[0286] In this orientation, the sub-frame motors 505 and fixed motors 504 generate lift in a direction that causes the aircraft to hover with the longitudinal axis of the fuselage 501A and wings 501B parallel to the ground (i.e., its forward flight orientation). In hover, the stability and controllability of the aircraft are controlled using the differential thrust between the fixed motors 504 and the total thrust of the sub-frame motors 505. An electronic flight controller regulates this differential thrust.

[0287] To initiate forward flight (FIG. 5), applying a relatively balanced, opposite differential thrust between the sub-frame motors 505 causes the sub-frame 502 to pitch forward, causing the aircraft 500 to begin forward flight. As speed increases, the wings 501B generate lift, keeping the aircraft 500 aloft. In some embodiments, reducing the thrust of the fixed motors 504 pitches the sub-frame 502 down until its thrust is directed forward.

[0288] During forward flight, the attitude of the aircraft is controlled by at least one of differential thrust between fixed motors 504, control surfaces (e.g., on wings 501B), and varying the angle and / or total thrust of subframe 502 as previously described.

[0289] Optionally, during forward flight, some control arrangement allows for stopping of stationary motor 504. The motor (per se, i.e., power source) of aircraft 500 optionally comprises at least one of the following power sources: electricity, internal combustion engine, and turbine. Air thrust propulsion of aircraft 500 optionally includes at least one of the following power sources: propeller, ducted fan, jet engine, and rocket booster.

[0290] (controllable pitch propeller) Reference is now made to Figures 7-9, which are schematic diagrams of electric motor arrangements according to some embodiments of the present disclosure, in which the angle of attack of propeller blades 705, 706 can be changed during propeller rotation by applying differential thrust between two parts of electric motor 700.

[0291] In some embodiments, varying the blade angle is used to optimize propeller efficiency and power output according to the axial speed of air through the propeller. Optionally, any of the aircraft embodiments described herein (e.g., the embodiments of FIGS. 1A-1O, 10A-11E, and 16) are provided with variable pitch blades.

[0292] The electric motor 700 (which is the "motor itself," i.e., the power source for moving the propellers 705A, 706A) comprises, in some embodiments, two stators 701, 702 having magnetic poles wound with coils of conductive wire for generating a magnetic field to rotate permanent magnet rotors 703, 704.

[0293] The two stators 701, 702 are fixed one in front of the other (the next two). The two stators 701, 702 may be similar or different in size and electrical properties.

[0294] Mounted around each stator 701, 702 is a corresponding free-spinning rotor 703, 704 on bearings or other couplings that allow rotation (it should be clarified that the rotors 703, 704 are electromechanical parts of the motor 700 itself, and are not self-contained propulsion units). Each rotor has a permanent magnet that faces the coils of the stator. The two rotors 703, 704 may have similar or different dimensions and magnetic properties.

[0295] In some embodiments, blades 705A, 705B (FIGS. 8 and 9) are attached to a single rotor 703 by a free-spinning hinge 707. The hinge allows the blades 705A, 705B to rotate in a direction that changes the blade angle of attack relative to the plane of rotation.

[0296] Each blade 705A is also connected to the second rotor 704, for example by a lever 705 hinged to the second rotor 704, or by a toothed mechanism 706, for example including a gear or part thereof having teeth that mesh with another gear or part thereof on the second rotor 704.

[0297] The connection to the second rotor 704 is configured so that a difference in position between the two rotors 703, 704 results in a change in the angle of attack of the blades 705A, 705B.

[0298] Power is provided from an electric motor control unit to each stator 701, 702. The output power of the two electric motor control units is commanded by an electronic controller.

[0299] The electronic controller receives throttle and propeller pitch commands, for example, from an aircraft controller and / or an electronic controller.

[0300] In some embodiments, the controller commands separate power levels to the two electronic motor controllers. Optionally, the sum of the power provided by the two units equals the thrust requested by the throttle command. Optionally, and without regard to any particular theory of operation, the difference in power between (i.e., applied to) the two electronic motor controllers (i.e., a change in power) causes an adjustment in the propeller pitch angle requested by the pitch command. Additionally or alternatively, and without regard to any particular theory of operation, changing the individual power levels commanded by the controllers changes the force applied to the base of the propeller blades, thereby adjusting the propeller pitch angle. Optionally, one of the electric motors is significantly more powerful than the other and provides the majority of the power that rotates the blades and generates thrust. The weaker electric motor is strong enough to overcome (leverage its connection to the blades) forces that tend to twist the propeller blades as they move through the air.

[0301] This variable pitch approach can also be used in other applications where an electric motor with a propeller is used to generate thrust in air or water.

[0302] Optionally, sensors are used to control and / or verify the pitch angle.

[0303] In some embodiments, each individual blade is provided with its own pitch control stator / rotor pair, e.g., a main power stator / rotor pair and four pitch control stator / rotor pairs, one for each of the four blades. The mechanical coupling is They are configured differently from one another. In some embodiments, this individualized blade attitude control is operated to continuously vary blade pitch during rotation to achieve cyclic blade control functions like those in a helicopter. A helicopter uses cyclic blade control to vary roll and pitch. While a multicopter can achieve roll / pitch control through differential actuation of its rotors, the addition of cyclic blade control allows, for example, rotors located closer to the fuselage (nearer the center of gravity) to provide greater attitude control than would otherwise be achieved.

[0304] (Optional electric motor design and control methods) Reference is now made to Figure 13, which is a schematic block diagram of an electric motor (per se) 1300, according to some embodiments of the present invention. Blocks in brackets indicate units of which the motor contains multiple units.

[0305] In some embodiments, electric motor 1300 (which may itself be, for example, electric motor 700 of FIG. 7 ) includes at least one stator (e.g., stator 1302) having coils 1311 of conductive wire, and rotor 1301 (i.e., the rotor component of electric motor 1300) including permanent magnets 1308, similar to a typical brushless electric motor. However, unlike a typical brushless electric motor, each coil 1311 on the stator of the electric motor embodiments described herein is wound with its own conductive wire and is not connected to the other coils 1311. In a typical brushless electric motor, several coils are wound with a single wire to form phases, and the phases are interconnected in different pair combinations (e.g., each of the six pair and polarity combinations in a three-phase motor) to continuously rotate the stator.

[0306] Also in some embodiments, each coil 1311 has its own control unit 1312, which includes power switching components that can optionally switch between supplying power to the coil with a given polarity or reverse polarity to generate a north or south magnetic field at the ends of the coil, or disconnecting power from the coil. Switching between these three options can be very fast, for example, up to several thousand hertz, and can be repeated.

[0307] The desired state of each switching unit of each coil 1311 is in turn determined by logic circuitry and / or a controlling microprocessor 1313 .

[0308] Optionally, a logic circuit or control microprocessor 1313 can be attached to each controller unit for each coil, or there can be a central logic circuit or control microprocessor (for example as shown) connected to and controlling the power switching units of all the coils.

[0309] Power from a battery or other external power source 1314 is distributed to all coil control units by a DC bus, of either positive or negative polarity, connected to all coil control units.

[0310] Optionally, a rotor position sensor 1315 (e.g., any of one or more types known from typical brushless motors) is used to determine the position of rotor 1301 relative to stator 1302, allowing logic circuitry or microprocessor 1314 to time switching during operation of motor 1300. Optionally, power is determined by sensing the current generated in a then-de-energized coil 1311 (using back-EMF).

[0311] A potential advantage of motor 1300 over a typical brushless motor is that it Unlike typical three-phase brushless motors, where only two-thirds of the coils are energized during a given period, all coils are used to power the motor most of the time, potentially increasing the torque and power per weight of the motor of the present invention relative to typical brushless motors.

[0312] In some embodiments, other potential advantages include reduced weight and / or power loss due to the elimination of wires connecting between the coils and phase wires exiting the motor to an external controller.

[0313] Another potential advantage is that while in a typical brushless electric motor, failure of the motor controller results in failure of the motor, electric motor 1300 includes redundancy such that failure of coil controller 1312 results in a fractional power degradation roughly proportional to the rated power of motor 1300 divided by the number of coils 1311.

[0314] There are also potential manufacturing advantages: whereas typical motor manufacturing processes require winding the entire motor, motor 1300 can be easily assembled from separate, identical coils 1311, each with its own control unit 1312. Also, during maintenance, it is possible to replace a single coil 1311 without having to unwind and rewind the entire motor.

[0315] Motor 1300 is optionally configured with stator 1302 as the inner part and permanent magnet rotor 1301 on the outside, or with stator 1302 on the outside and permanent magnet rotor 1301 on the inside, or with stator 1302 and rotor 1301 stacked one on top of the other in a direction along the axis of rotation of motor 1300, or any other relative position used in a typical brushless motor.

[0316] Any size and number of stator coil 1311 and rotor magnet 1308 pairs can be used.

[0317] (propeller guard) Reference is now made to Figures 14A-14C, which schematically illustrate an angled propeller guard 1401 for an aircraft, according to some embodiments of the present disclosure.

[0318] 14A shows a rotor device 1400 including a rotor guard 1401 in conjunction with a rotor 1402. The rotor 1402 includes at least one motor 1406 and associated propeller blades 1405. In some embodiments, the rotor guard 1401 is oriented generally in the same plane as the propeller 1402.

[0319] Figure 14B shows rotor guard 1401 as seen from the side of the aircraft. Figure 14C shows rotor guard 1401 as seen from the front of the aircraft, and slightly above rotor guard 1401. In some embodiments, rotor guard 1401 has a forward portion 1401A that is horizontally (radially) long and vertically short, with a relatively flattened cross-section. Optionally, this approximates an airfoil or wing-like cross-section. Optionally, this airfoil-like cross-section is oriented at an optimum angle of attack similar to the optimum angle of attack of aircraft wing 120.

[0320] In some embodiments, the central portion 1401B of the rotor guard 1401 is more vertically oriented, i.e., longer vertically and shorter horizontally (in the horizontal radial direction). The aft portion 1401C is optionally flattened in cross section, e.g., of the same shape as the forward portion 1401A. The overall shape of the rotor guard 1401 may optionally be understood as consisting generally of a short, oblique cylinder, with the top and bottom of the short cylinder offset in one direction relative to each other. Optionally, Optionally, the offset allows the rotor guard 1401 to have a minimal profile relative to the forward flight direction.

[0321] Potential benefits of the rotor guard 1401 include drag minimization by the propeller guard 1401, as well as potential lift contributions from the propeller guard 1401. In some embodiments, the overall rotor device 1400 includes multiple rotors (e.g., a stack of contra-rotating propellers). The rotor guard 1401 is optionally angled to intersect the propeller stack at a small oblique angle, for example, flush with the lower rotor at the front and flush with the upper rotor at the rear.

[0322] (Cooperative power supply) Reference is now made to FIG. 15 , which illustrates a schematic diagram of power connections between multiple battery units 1501A, 1501B, 1501C, and 1501D and multiple motors 1502A, 1502B, 1502C, 1502D, 1503A, 1503B, 1503C, and 1503D, according to some embodiments of the present disclosure. The motors are shown approximately as physically located around the aircraft. The physical placement of the battery units is optionally in any suitable arrangement. The motors are paired as counter-rotating, double-rotor assemblies, i.e., 1502A and 1503A, 1502B and 1503B, 1502C and 1503C, and 1502D and 1503D.

[0323] Several embodiments of the present disclosure describe motor arrangements that include motors located at diagonally opposite corners of the aircraft. One potentially unstable flight mode results from one or more rotors at one corner of the aircraft becoming weak or inoperable. In this mode, one or more rotors at the opposite corner may provide enough thrust to flip the aircraft before controls can be exercised to prevent this. For example, if motor 1502B becomes inoperative, any thrust applied by motor 1502C will attempt to rotate the entire aircraft about an axis extending approximately parallel to the line connecting motors 1502A and 1502D.

[0324] In an electrically powered system, one way rotor failure can occur is through battery failure. In some embodiments of the present disclosure, rotor power is distributed among multiple battery units so that a battery unit failure affects pairs of rotors that have a mutual balancing role in stabilizing the aircraft. Optionally (e.g., as shown in FIG. 15), multiple separate battery units power each pair of rotors.

[0325] Thus, for example, motors 1502B and 1502C are commonly connected to battery unit 1501A, and motors 1503B and 1503B are commonly connected to battery unit 1501D. Motors 1502A and 1502D are commonly connected to battery unit 1501B, and motors 1503A and 1503D are commonly connected to battery unit 1501C. If any one battery unit fails, the remaining rotor set can still be balanced and operational. For purposes of this description, rotors in a "balancing" rotor set are positioned to act from opposite corners with respect to a frame of reference established perpendicular to the aircraft's center of gravity and the direction of forward flight motion. That is, they are positioned laterally offset from the center of gravity, away from both the central axis extending from the front to the rear of the aircraft and a horizontal axis intersecting the central axis. In some embodiments, the balancing rotors further include at least one rotor positioned therebetween. For example, it would otherwise tend to extend an uncontrolled diagonal axis of rotation between them.

[0326] (Flight Controller) Reference is now made to FIG. 17, which schematically illustrates a control unit 1700 of a distributed flight control system, according to some embodiments of the present disclosure.

[0327] In some embodiments, all elements of control unit 1700 are provided for each individual rotor (and, optionally, for each motor / propeller subunit of a multi-propeller rotor), which offers the potential advantage that failure of one flight controller 1701 only significantly affects the operation of one motor.

[0328] In some embodiments, flight controller 1701 includes at least one of an optional inertial measurement unit (IMU) 1702, a module 1703 that performs velocity calculations (in each of the three spatial axes), and a module 1704 that performs attitude calculations based on flight metrology data 1714 received from IMU 1702, flight metrology data 1711 received from other flight controllers 1701, and optionally flight metrology data 1712 from a central IMU. In embodiments with a separate IMU per control unit 1700, flight metrology data 1713 is output to the other flight controllers.

[0329] The current calculated velocity 1715 and attitude 1716 are forwarded to command computation unit 1705, which uses them along with control input 1718 (e.g., received from at least one of a control stick, a flight controller, and an attitude command) to generate output commands 1710 that are provided to motor controller 1706. Motor controller 1706 then controls the power of motor 1707 (e.g., the rotor's power source). Optionally, the control commands sent with output commands 1710 include other aspects of motor 1707 operation, such as blade pitch, differential control of propellers in a multi-propeller rotor, or another control aspect.

[0330] In some embodiments, each flight controller 1701 is located near a motor 1707. As a result, IMU 1702 sensing automatically incorporates fuselage and / or strut flex (aircraft frame flex), potentially improving control response time and / or accuracy. For example, when a motor command is given, the complete or immediate change in rotor thrust direction may not be sensed by the central IMU because mounting strut flex absorbs a portion of the rotor thrust attitude change. Locating IMU 1702 near motor 1707 may reduce this type of sensing distortion.

[0331] Optionally, an average or other combination and / or selection of sensor data is used. In some embodiments, the target state of the aircraft is based on a current estimated flight state that is the same for all flight controllers 1701 and accessible by all flight controllers 1701. In some embodiments, the estimated current flight state is calculated by, for example, a central controller that receives inputs from and distributes results back to multiple flight controllers of the aircraft. Additionally or alternatively, the estimated current flight state for the entire aircraft is calculated at each individual flight controller 1701 using flight data 1711 from the same common group of IMUs 1702, optionally from all IMUs 1702. Optionally, data from one or more IMUs is excluded from the calculation, for example, because it cannot be correlated with flight data from other IMUs 1702 or for another reason.

[0332] In some embodiments, local IMU data 1714 is prioritized in velocity and attitude calculations, but is validated using comparison with other available IMU data 1711, 1712. In some embodiments, data 1714 from the local IMU is given special priority and / or relevance during more dynamic periods of flight (e.g., changing control inputs), which may increase response speed and reduce control resonance.

[0333] With regard to attitude in particular, the attitude calculation module 1704 optionally uses an average of all available inputs or other combination and / or selection (e.g., the most extreme sensed attitude consistent with normally expected measurement variations) to potentially avoid distortions in attitude sensing that would cause different rotors to steer to different attitudes.

[0334] If local data shows unreasonable discrepancies from the overall agreement of other inputs (e.g., if it differs from most other inputs by a physically unrealistic amount, assuming the aircraft is undamaged), it may be optionally ignored in favor of other available data. Similarly, individual inputs from other data sources may optionally be ignored if they give unreasonable readings, for example, that contradict most other data sources.

[0335] Optionally, other available IMU data 1711, 1712 allows the local flight controller 1701 to make more advanced control decisions, for example, taking into account how other control units 1700 are likely to react. Optionally, the rotors are configured to provide "watchdog" signals to each other to indicate their continued functional status and to report to each other when they fail to operate normally or partially. Optionally, the flight control system module enters a fallback mode, relying more fully on independent control (or alternatively, more fully on centralized control), when it detects that one or more rotors have stopped reporting and / or operating normally. Optionally, the choice between centralized and independent control depends on the pattern of failures detected.

[0336] However, there are potential advantages in terms of simplicity and / or analyzability of flight characteristics for each flight controller to, at least typically, rely entirely or almost entirely on its own sensor readings, which in either case can also be understood as essentially "sensing" the composite behavior of the rest of the control system collectively. As long as each individual flight controller 1701 "knows" what the state of its own IMU data should be in response to a given set of control inputs 1718, it can act to increase power, decrease power, or otherwise control motors 1707 to move the aircraft at speed and direction toward its target state.

[0337] In some embodiments of the invention, flight controller 1701 (optionally another flight controller configuration) is configured to control yaw by controlling the rotational speed of only one propeller of a rotor with multiple counter-rotating propellers. Optionally, the rotor is oriented at a tilt that directs some thrust in the yaw direction. This control method potentially allows both yaw thrust and yaw torque to be used, for example, to exercise yaw authority as described in the overview.

[0338] (General remarks) It is anticipated that many related power sources for driving thrust-producing motors will be developed during the life of the patent expiring from this application, and the scope of the term power source is intended to include such new technology a priori.

[0339] The term "about" as used herein in connection with an amount or value means "within ±10%."

[0340] The terms "comprises," "comprising," "includes," "including," "having," and their conjugations mean "including but not limited to."

[0341] The term "consisting of" means "including and limited to."

[0342] The term "consisting essentially of" means that a composition, method, or structure may include additional components, steps, and / or portions, but only if the additional components, steps, and / or portions do not materially alter the basic and novel characteristics of the claimed composition, method, or structure.

[0343] As used herein, the singular forms "a," "an," and "the" include plural references unless the context clearly dictates otherwise. For example, the term "a compound" or "at least one compound" can include multiple compounds, including mixtures thereof.

[0344] The words "example" and "exemplary" are used herein to mean "serving as an example, instance, or illustration." Any embodiment described as "example" or "exemplary" is not necessarily to be construed as preferred or advantageous over other embodiments, or as excluding the incorporation of features of other embodiments, or both.

[0345] The term "optionally" is used herein to mean "provided in some embodiments and not provided in other embodiments." Any particular embodiment of the present disclosure may include multiple "optional" features unless such features are contradictory.

[0346] As used herein, the term "method" refers to ways, means, techniques, and procedures for accomplishing a given task, including, but not limited to, such ways, means, techniques, and procedures that are known to or readily developed by those skilled in the art of chemistry, pharmacology, biology, biochemistry, and medicine from known ways, means, techniques, and procedures.

[0347] As used herein, the term "treating" includes negating, substantially inhibiting, slowing or reversing the progression of a condition, substantially ameliorating the clinical or cosmetic symptoms of a condition, or substantially preventing the appearance of clinical or cosmetic symptoms of a condition.

[0348] Throughout this application, embodiments may be presented with reference to a range format. It should be understood that the description in range format is merely for convenience and simplicity and should not be construed as a fixed limitation on the scope of the descriptions in this disclosure. Accordingly, descriptions of ranges should be considered to specifically disclose all possible subranges along with individual numerical values ​​within that range. For example, a description of a range such as "1 to 6" should be considered to have specifically disclosed subranges such as "1 to 3," "1 to 4," "1 to 5," "2 to 4," "2 to 6," "3 to 6," etc., along with individual numerical values ​​within that range, e.g., 1, 2, 3, 4, 5, and 6. This applies regardless of the breadth of the range.

[0349] Whenever a numerical range is given herein (e.g., "10-15," "10 to 15," or any pair of numbers joined by another such range indicator), it is meant to include any number (fractional or integer) within the limits of the stated range, inclusive of the limits of that range, unless the context clearly dictates otherwise. The phrases "ranging between / across / between" a first specified number and a second specified number, and the phrases "from" a first specified number to a second specified number, "up to," "until," "through" and "including" a second specified number, are used interchangeably herein and mean to include the first specified number and the second specified number, and all fractional and integer numbers therebetween.

[0350] While the description of this disclosure has been provided in conjunction with specific embodiments, it is evident that many alternatives, modifications, and variations will be apparent to those skilled in the art. Accordingly, the description of the disclosure is intended to embrace all such alternatives, modifications, and variations that fall within the spirit and broad scope of the appended claims.

[0351] All publications, patents, and patent applications mentioned in this specification are herein incorporated by reference in their entirety to the same extent as if each individual publication, patent, or patent application was specifically and individually indicated to be incorporated by reference. Furthermore, citation or identification of any reference in this application shall not be construed as an admission that such reference is available as prior art to the present disclosure. To the extent used as section headings, they should not be construed as necessarily limiting.

[0352] It is understood that certain features, which for clarity are described in this disclosure in the context of separate embodiments, may also be provided in combination in a single embodiment. Conversely, various features of the invention that are for brevity described in the context of a single embodiment may also be provided separately or in any suitable subcombination, or as appropriate with any other described embodiment of this disclosure. Certain features described in the context of various embodiments are not considered essential features of those embodiments, unless the embodiment is inoperable without those elements.

[0353] Additionally, any priority documents of this application are incorporated herein by reference in their entirety.

Claims

1. 1. A multi-rotor aircraft, comprising: The torso and a plurality of rotors connected to the fuselage; a plurality of independently operating flight controller units, each flight controller unit including an inertial measurement unit (IMU); Equipped with each of the rotors is individually controlled by a respective flight controller unit based on measurements from a respective IMU of the flight controller unit; Multirotor aircraft.

2. The multi-rotor aircraft of claim 1 , wherein each of the plurality of flight controller units receives measurements from a respective IMU of the other flight controller units.

3. 3. The multi-rotor aircraft of claim 1 or claim 2, wherein the multiple flight controller units share IMU data, with each flight controller unit having access to an estimate of overall aircraft flight conditions.

4. 4. The multi-rotor aircraft of claim 3, wherein each flight controller unit issues commands intended to result in the same overall aircraft flight state based on an estimate of the overall aircraft flight state.

5. 5. The multi-rotor aircraft of claim 1, wherein each flight controller unit is configured to issue commands to a description of overall aircraft flight conditions modified by changes in its IMU measurements, the modifications including adjustments in response to flexing of at least one of struts and a frame of the multi-rotor aircraft.

6. 6. The multi-rotor aircraft of claim 1, wherein each of the plurality of flight controller units is configured to issue commands that take into account how other flight controller units are likely to react.

7. The multi-rotor aircraft of any one of claims 1 to 6, wherein the rotors are configured to provide each other with watchdog signals indicative of their continued functioning status.

8. The multi-rotor aircraft of any one of claims 1 to 6, wherein each flight controller unit comprises a speed calculation module that performs speed calculations based on flight measurement data received from its respective IMU.

9. The multi-rotor aircraft of claim 8 , wherein each flight controller unit includes an attitude calculation module that performs attitude calculations based on flight measurement data received from the respective IMU.

10. 10. The multi-rotor aircraft of claim 9, wherein the velocity calculation module and the attitude calculation module are configured to prioritize data from the respective IMUs and to validate the data from the respective IMUs with comparison to other available data from IMUs of other flight controller units.

11. 10. The multi-rotor aircraft of claim 9, wherein the velocity calculation module and the attitude calculation module are configured to prioritize the data of the respective IMUs during more dynamic periods of flight.

12. 10. The multi-rotor aircraft of claim 9, wherein the velocity calculation module and the attitude calculation module are configured to ignore input from the respective IMU or input from the other IMU if the data from the respective IMU or the data from the other IMU exhibits an unreasonable discrepancy from a general agreement of data from the other IMU.

13. The multi-rotor aircraft of claim 9 , comprising at least one of a central controller and a central IMU.

14. 14. The multi-rotor aircraft of claim 13, wherein the central controller receives inputs from the plurality of flight controller units, calculates estimates of overall aircraft flight states, and distributes the estimates of overall aircraft flight states back to the plurality of flight controller units.

15. 15. The multi-rotor aircraft of claim 14, wherein the velocity calculation module and the attitude calculation module are configured to enter a fallback mode relying entirely on the central controller if they detect that one or more rotors have stopped reporting or operating normally.

16. The multi-rotor aircraft of claim 15 , wherein the velocity calculation module and the attitude calculation module are configured to select reliance on the central controller depending on a pattern of detected faults.

17. A multi-rotor aircraft according to any one of claims 1 to 16, a plurality of motor assemblies, each configured to generate thrust by moving air past the motor assembly along a respective thrust axis of the motor assembly, each motor assembly comprising one of the plurality of rotors; Wings and Equipped with During operation of the multi-rotor aircraft, the direction of each thrust axis is fixed at a constant pitch angle inclined to the pitch direction of the wing; the plurality of motor assemblies together are operable to both fully support the multi-rotor air vehicle in a hover mode and to forwardly propel the multi-rotor air vehicle in a forward flight mode; each motor assembly is positioned such that the wings do not cross, within a radius of the motor assembly, airflow before or after entering the motor assembly along each thrust axis; Multirotor aircraft.

18. 1. A method of controlling a multi-rotor aircraft, the multi-rotor aircraft comprising: The torso and a plurality of rotors connected to the fuselage; a plurality of independently operating flight controller units, each flight controller unit comprising an inertial measurement unit (IMU); wherein the method comprises: measuring flight measurement data by each IMU of each flight controller unit; Independently controlling each of the rotors based on measurements from each IMU of the flight controller unit.

1. A method for controlling a multi-rotor aircraft, comprising:

19. and further comprising operating one flight controller unit of the plurality of flight controller units, said operating comprising: receiving flight measurement data from an IMU of the flight controller unit; a speed calculation module performs speed calculation based on the flight measurement data; an attitude calculation module performs attitude calculation based on the flight measurement data; generating output commands based on the velocity calculations and the attitude calculations; providing the output command to a motor controller; 20. The method of controlling a multi-rotor aircraft of claim 18, comprising:

20. 20. The method of controlling a multi-rotor aircraft of claim 18 or claim 19, further comprising outputting the flight measurement data to another flight controller unit.

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