Hybrid aerial vehicle with multiple flight modes

Abstract

A hybrid aerial vehicle with multiple flight modes is provided, including: a central member; a first wing member arranged at a first side of the central member, the first wing member having a first thrust unit and a first flap actuator affixed thereto, the first flap actuator having a first flap member affixed thereto; a second wing member arranged at a second side of the central member, the second wing member having a second thrust unit and a second flap actuator affixed thereto, the second flap actuator having a second flap member affixed thereto; a wing actuator configured as a linear actuator and configured to rotate the first and second wing members about a lateral axis of the central member via a first mechanical linkage and a second mechanical linkage, respectively; and a flight controller communicatively coupled to the wing actuator, the first and second thrust units and the first and second flap actuators and configured to control, in response to a transition control signal received, the hybrid aerial vehicle to transition from one flight mode to another flight mode amongst the multiple flight modes during flight. In particular, the multiple flight modes include a rotating rotorcraft flight mode, a non-rotating rotorcraft flight mode and a fixed-wing flight mode.

Description

HYBRID AERIAL VEHICLE WITH MULTIPLE FLIGHT MODESCROSS-REFERENCE TO RELATED APPLICATION

[0001] This application claims the benefit of priority of Singapore Patent Application No. 10202251650R, filed on 7 November 2022, the content of which being hereby incorporated by reference in its entirety for all purposes.TECHNICAL FIELD

[0002] The present invention generally relates to a hybrid aerial vehicle with multiple flight modes, a method of operating the hybrid aerial vehicle and a method of forming the hybrid aerial vehicle.BACKGROUND

[0003] The concept of an aerial vehicle (e.g., Unmanned Aerial Vehicle (UAV)) with multiple modes of operation (multiple flight modes) has proven advantageous for aerial vehicles to traverse different flight envelopes, bringing more versatility and flexibility to modem aerial vehicle (e.g., UAV) operations. For example, in a first responder application, the ability to have a quick response time across long distances and also the ability to hover at any point in time (e.g., over rivers, mountains or urban environments) is an added advantage. Another such application is in surveillance / security applications, where having a UAV that can rapidly deploy and traverse large distances paired with the ability to loiter in the air is a huge asset. Accordingly, a hybrid system that can dynamically switch its mode of operation (flight mode) in mid-air can access different advantages that each distinct flight mode brings depending on the current operational requirement at hand. For example, a system operating as a rotorcraft gains maneuverability and the ability to perform a stationary hover, while sacrificing range and speed associated with a fixed-wing aircraft. On the other hand, for example, a system operating as a fixed-wing aircraft gains the longer endurance and range that a rotorcraft typically does not possess, however, it needs to maintain a forward velocity to remain in the air.

[0004] There have been disclosed hybrid aerial vehicles having two flight modes. For example, there has been disclosed a hybrid UAV where a quadrotor frame is simply fused with a fixed-wing frame. Although such a configuration may be simple, it is not very structurally efficient due to redundant components. For example, there has also been disclosed a hybrid UAV known as the T-Wing, which demonstrated the feasibility of tail sitter-type aircraft wherethe control surfaces are submerged in the propeller wash of the propulsion system. Some other previous works include a tilt-rotor UAV design that actuates the angle of the propulsion system to accommodate the different flight modes. For example, there has been disclosed a quadrotor tail sitter for autonomous parcel delivery. This quadrotor tail sitter has four motors with no wing control surfaces, and is controlled purely by differential thrust. For example, commercially available hybrid UAV designs include the FireFly6 by BirdsEyeView Aerobotics and WingtraOne by Wingtra. However, these hybrid aerial vehicles with two flight modes mostly draw their efficiency from the fixed-wing flight mode, and only use the hover mode as a means for vertical take-off and landing.

[0005] A need therefore exists to provide a hybrid aerial vehicle (which may also be referred to as an aircraft or an aerial robot) with multiple flight modes, that seeks to overcome, or at least ameliorate, one or more deficiencies in conventional hybrid aerial vehicles, and more particularly, provide a hybrid aerial vehicle with multiple flight modes with improved operational efficiency and effectiveness. It is against this background that the present invention has been developed.SUMMARY

[0006] According to a first aspect of the present invention, there is provided a hybrid aerial vehicle with multiple flight modes, comprising: a central member; a first wing member arranged at a first side of the central member, the first wing member having a first thrust unit and a first flap actuator affixed thereto, the first flap actuator having a first flap member affixed thereto and is configured to actuate the first flap member; a second wing member arranged at a second side of the central member, the second wing member having a second thrust unit and a second flap actuator affixed thereto, the second flap actuator having a second flap member affixed thereto and is configured to actuate the second flap member, wherein the first and second sides of the central member are opposite sides of the central member; a wing actuator configured as a linear actuator and configured to rotate the first and second wing members about a lateral axis of the central member via a first mechanical linkage and a second mechanical linkage, respectively; and a flight controller communicatively coupled to the wing actuator, the first and second thrust units and the first and second flap actuators and configured to control, in response to atransition control signal received, the hybrid aerial vehicle to transition from one flight mode to another flight mode amongst the multiple flight modes during flight based on controlling the wing actuator, the first and second thrust units and the first and second flap actuators, the multiple flight modes comprising a rotating rotorcraft flight mode, a non-rotating rotorcraft flight mode and a fixed-wing flight mode.

[0007] According to a second aspect of the present invention, there is provided a method of operating the hybrid aerial vehicle according to the above-mentioned first aspect of the present invention, the method comprising: controlling, by the wing controller and in response to the transition control signal being for transitioning the hybrid aerial vehicle from the rotating rotorcraft flight mode to the nonrotating rotorcraft flight mode, the wing actuator to rotate the first and second wing members about the lateral axis of the central member from the extended state to the retracted state, and controlling, by the wing controller and in response to the transition control signal being for transitioning the hybrid aerial vehicle from the non-rotating rotorcraft flight mode to the rotating rotorcraft flight mode, the wing actuator to rotate the first and second wing members about the lateral axis of the central member from the retracted state to the extended state.

[0008] According to a third aspect of the present invention, there is provided a method of forming the hybrid aerial vehicle according to the above-mentioned first aspect of the present invention.BRIEF DESCRIPTION OF THE DRAWINGS

[0009] Embodiments of the present invention will be better understood and readily apparent to one of ordinary skill in the art from the following written description, by way of example only, and in conjunction with the drawings, in which:FIG. 1 depicts a schematic drawing of a hybrid aerial vehicle with multiple flight modes, according to various embodiments of the present invention;FIG. 2 depicts a schematic flow diagram of a method of operating the hybrid aerial vehicle, according to various embodiments of the present invention;FIG. 3 depicts a schematic block diagram of the flight controller configured to control, in response to a transition control signal received, the hybrid aerial vehicle to transition from one flight mode to another flight mode amongst the multiple flight modes during flight, according to various embodiments of the present invention;FIG. 4 depicts a schematic flow diagram of a method of forming the hybrid aerial vehicle with multiple flight modes, as described hereinbefore according to various embodiments the present invention;FIG. 5A shows example maple seeds;FIG. 5B shows a composite image of the autorotation motion of a maple seed;FIG. 5C depicts a multi-modal hybrid aerial vehicle under the rotor-wing mode (which corresponds to the rotating rotorcraft flight mode and may be referred to as the R-mode) according to various example embodiments of the present invention;FIGs. 6A to 6C show an overview of the three flight modes of the hybrid UAV, namely, rotating rotorcraft flight mode (which may be referred to as the R-mode), non-rotating rotorcraft flight mode (which may be referred to as the T-mode) and fixed-wing flight mode (which may be referred to as the C-mode), respectively, according to various example embodiments of the present invention;FIG. 7 depicts a schematic drawing of an example hybrid UAV with the three flight modes, according to various embodiments of the present invention;FIG. 8 depicts the UAV in each of the three flight modes and transitions therebetween, according to various example embodiments of the present invention;FIG. 9 A shows the UAV in the R-mode and the T-mode, along with the inertial frame J and the UAV body frame B, according to various example embodiments of the present invention;FIG. 9B illustrates aerodynamic forces on a single blade element, according to various example embodiments of the present invention;FIG. 10A depicts a cross-sectional front view of the UAV (corresponding to the T- or C-mode position), excluding the wings for visual clarity, according to various example embodiments of the present invention;FIG. 10B depicts a schematic drawing of the mechanical linkage showing the relations between various components thereof when the linear actuator is in an extended state during hover in the R-mode and when the linear actuator is in a retracted state during the T- or C-mode, according to various example embodiments of the present invention;FIG. IOC depicts a schematic drawing of a cross-section of the mechanical linkage in the extended state, according to various example embodiments of the present invention;FIG. 11 A depicts a schematic block diagram of an example flight controller of the UAV, according to various example embodiments of the present invention;FIG. 11B depicts a schematic block diagram of the R-mode attitude controller, according to various example embodiments of the present invention;FIG. 11C illustrates the tip path plane (TPP) of the UAV in the R-mode, according to various example embodiments of the present invention;FIG. 1 ID depicts a schematic drawing of an example central pattern generator (CPG) controller of the R-mode attitude controller, according to various example embodiments of the present invention;FIG. HE depicts a schematic block diagram of the T / C-mode attitude controller, according to various example embodiments of the present invention;FIGs. 12A and 12B depict a plot of control weights (sigmoid control weights) for the forward transition direction and the backward transition direction, respectively, and illustrates the effect of various function parameters on the shape of the control weights, according to various example embodiments of the present invention;FIG. 13 shows the blending of control output commands from the R-mode attitude controller and the T / C-mode attitude controller to produce the final control output commands for controlling the flight of the UAV, based on the angle of the wings (Bangieand the fixed- wing transition state (Bcruise), according to various example embodiments of the present invention;FIG. 14 shows a table (Table 1) of example blending parameters which may be optimized, according to various example embodiments of the present invention;FIG. 15 shows a table (Table 2) of example constraints of the blending parameters and their respective penalty functions, according to various example embodiments of the present invention;FIG. 16 depicts a schematic flow diagram of the optimization process associated with each episode to obtain the objective function, according to various example embodiments of the present invention;FIG. 17 shows a table (Table 3) of the solution for the parameter values returned by the optimization process, according to various example embodiments of the present invention;FIG. 18 depicts plots of the values of the control weights (fe, br, be) in relation to the transition coordinator variables (Bangle, Bcruise for a switched controller (top) and the optimized blended controller (bottom), according to various example embodiments of the present invention;FIG. 19 depicts a plot of the fitness value throughout the optimization, according to various example embodiments of the present invention;FIG. 20A shows plots of the XY position histories of 10 consecutive simulations from R-T modes, which present the transition simulation data for unoptimized switched control transitions and optimized blended control transitions;FIG. 20B shows the simulated throttle comparison of the R-mode to the T-mode transition (R-to-T transition) and of the C-mode to R-mode transition (C-T-R transition) for the unoptimized switched control and optimized blended transitions;FIGs. 21A and 21B depict photos of a UAV prototype used in experiments conducted in the R-mode and in the T / C-mode, respectively, according to various example embodiments of the present invention;FIG. 22 depicts a schematic block diagram of various electronics components onboard the UAV prototype and offboard, as well as example communication paths (coupling / connections) between various electronics components, according to various example embodiments of the present invention;FIG. 23 shows a table (Table 4) of various physical properties of the prototype UAV, according to various example embodiments of the present invention;FIG. 24A shows a plot of the angular velocities in the body frame during the repeated R-T-R transitions which were invoked manually by a user, according to various example embodiments of the present invention;FIG. 24B shows plots of the altitude data of the R-to-T transitions and T-to-R transitions of the three transition controls, according to various example embodiments of the present invention;FIG. 24C shows plots of the summed throttle data of the R-to-T transitions and T-to-R transitions of the three transition controllers, according to various example embodiments of the present invention;FIG. 25 shows a table (Table 5) of the Altitude Mean Absolute Error (MAE) and Throttle Sum (normalized) during R-to-T mode transitions as well as T-to-R mode transitions, according to various example embodiments of the present invention;FIG. 26A depicts plots of mean XY position history for each of the three different transition controls being compared, where the desired direction to transition to the C-mode is in the positive Y-axis direction;FIG. 26B depicts plots of the throttle usage comparisons of the three different transition controls;FIG. 27A shows a plot of the UAV position history in a map overlay of the test site for a flight;FIG. 27B shows a plot of the UAV position (in meters) against time and a plot of the UAV altitude against time;FIG. 28A shows a composite top-down aerial view of the transition process captured from another drone, which depicts the motion of the UAV prototype as it transitions between the three different flight modes (T-mode, R-mode and C-mode), as well as snapshots of the UAV prototype during the transition, according to various example embodiments of the present invention;FIG. 28B shows the position history of the UAV prototype during a hairpin-like maneuver, according to various example embodiments of the present invention; andFIG. 28C shows a boxplot of the power consumption during each of the three flight modes measured from onboard current and voltage sensors, according to various example embodiments of the present invention.DETAILED DESCRIPTION

[0010] Various embodiments of the present invention provide a hybrid aerial vehicle with multiple flight modes, a method of operating the hybrid aerial vehicle and a method of forming the hybrid aerial vehicle.

[0011] As discussed in the background, an aerial vehicle (e.g., UAV) with multiple modes of operation (multiple flight modes) (which may herein be referred to as a hybrid aerial vehicle) has proven advantageous for aerial vehicles to traverse different flight envelopes, bringing more versatility and flexibility to modern aerial vehicle (e.g., UAV) operations. In particular, a hybrid aerial vehicle that can dynamically switch its mode of operation (flight mode) in mid-air can access different advantages that each distinct flight mode brings depending on the current operational requirement at hand. There have been disclosed hybrid aerial vehicles having two flight modes. However, as discussed in the background, these hybrid aerial vehicles with two flight modes mostly draw their efficiency from the fixed-wing flight mode, and only use the hover mode as a means for vertical take-off and landing. In this regard, various embodiments provide a hybrid aerial vehicle (which may also be referred to as an aircraft or an aerial robot) with multiple flight modes, that seeks to overcome, or at least ameliorate, one or moredeficiencies in conventional hybrid aerial vehicles, and more particularly, provide a hybrid aerial vehicle with multiple flight modes with improved operational efficiency and effectiveness.

[0012] FIG. 1 depicts a schematic drawing of a hybrid aerial vehicle 100 with multiple flight modes according to various embodiments of the present invention. The hybrid aerial vehicle 100 comprises: a central member 110; a first wing member 120a arranged at a first side of the central member 110, the first wing member 120a having a first thrust unit 124a and a first flap actuator 126a affixed thereto, the first flap actuator 126a having a first flap member 128a affixed thereto and is configured to actuate the first flap member 128a (e.g., to adjust the flap angle / deflection); a second wing member 120b arranged at a second side of the central member 110, the second wing member 120b having a second thrust unit 124b and a second flap actuator 126b affixed thereto, the second flap actuator 126b having a second flap member 128b affixed thereto and is configured to actuate the second flap member 128b (e.g., to adjust the flap angle / deflection), whereby the first and second sides of the central member 110 are opposite sides of the central member 110; a wing actuator 130 configured as a linear actuator and configured to rotate the first and second wing members 120a, 120b about a lateral axis 140 of the central member 110 via a first mechanical linkage 150a and a second mechanical linkage 150b, respectively; and a flight controller 160 communicatively coupled to the wing actuator 130, the first and second thrust units 124a, 124b and the first and second flap actuators 126a, 126b and configured to control, in response to a transition control signal received, the hybrid aerial vehicle 100 to transition from one flight mode to another flight mode amongst the multiple flight modes during flight based on controlling the wing actuator 130, the first and second thrust units 124a, 124b and the first and second flap actuators 126a, 126b. In particular, the multiple flight modes comprise a rotating rotorcraft flight mode, a non-rotating rotorcraft flight mode and a fixed-wing flight mode.

[0013] In various embodiments, the rotating rotorcraft flight mode is a flight mode whereby the aerial vehicle is operating as a rotating rotorcraft and the flight of the aerial vehicle (e.g., lift thereof) is primarily achieved by rotating / spinning the entire aerial vehicle (with its wing members (e.g., rotating blades)) about its yaw axis through thrust generated by the thrust units (e.g., motor-driven propellers). The non-rotating rotorcraft flight mode is a flight mode whereby the aerial vehicle is operating as a non-rotating rotorcraft and the flight of the aerial vehicle (e.g., lift thereof) is primarily achieved by thrust units (e.g., motor-driven propellers) without necessarily rotating / spinning the entire aerial vehicle about its yaw axis. The fixed-wing flightmode is a flight mode whereby the aerial vehicle is operating as a fixed-wing aircraft and the flight of the aerial vehicle (e.g., lift thereof) is primarily achieved from the wings by maintaining a forward velocity through thrust generated by thrust units.

[0014] Accordingly, the hybrid aerial vehicle 100 according to various embodiments of the present invention is advantageously configured with multiple flight modes with improved operational efficiency and effectiveness. In particular, in contrast to conventional hybrid aerial vehicles with only two flight modes which, for example, mostly draw their efficiency from the fixed-wing flight mode, the hybrid aerial vehicle 100 is advantageously configured with at least two efficiency regions (at least efficient regions in hovering and cruising) and yet do not compromise on maneuverability by additionally providing a third flight mode, namely, the nonrotating rotorcraft flight mode for hovering. This is such that in addition to an efficient cruising state under the fixed-wing flight mode, an efficient hovering state can be achieved under the rotating rotorcraft flight mode as desired or required while enabling the flexibility to enter into an agile hovering state under the non-rotating rotorcraft flight mode (more agile but less efficient than the rotating rotocraft flight mode). Furthermore, not only is the hybrid aerial vehicle 100 capable of multiple flight modes, the hybrid aerial vehicle 100 comprises a flight controller 160 configured to control, in response to a transition control signal received, the hybrid aerial vehicle 100 to transition from one flight mode to another flight mode amongst the multiple flight modes during flight in an effective manner, and more particularly, based on controlling the wing actuator 130, the first and second thrust units 124a, 124b and the first and second flap actuators 126a, 126b during a transition phase in a collective or integrated manner. Accordingly, the hybrid aerial vehicle 100 according to various embodiments of the present invention is advantageously configured with multiple flight modes with improved operational efficiency and effectiveness. These advantages or technical effects, and / or other advantages or technical effects, will become more apparent to a person skilled in the art as the hybrid aerial vehicle 100 is described in more detail according to various embodiments and example embodiments of the present invention.

[0015] In various embodiments, the first and second mechanical linkages 150a, 150b (e.g., at respective first ends thereof) are coupled to the first and second wing members 120a, 120b, respectively, and are each (e.g., at a respective second end thereof, opposite to the respective first end) further coupled to the wing actuator 130. In various embodiments, the first and second mechanical linkages 150a, 150b are each configured to transform a linear motion of the wing actuator 130 (e.g., at the respective first end of the first / second mechanical linkage 150a / 150b)to a rotational motion (e.g., at the respective second end of the first / second mechanical linkage 150a / 150b) for rotating the corresponding wing member of the first and second wing members 120a, 120b about the lateral axis 140 of the central member 110.

[0016] In various embodiments, the wing actuator 130 is configured to rotate the first and second wing members 120a, 120b about the lateral axis 140 of the central member 110 in opposite directions (e.g., clockwise and anti-clockwise directions) via the first and second mechanical linkages 150a, 150b, respectively.

[0017] In various embodiments, the first thrust unit 124a and the first flap member 128a are arranged at a leading edge and a trailing edge of the first wing member 120a, respectively. For example, the first thrust unit 124a may be arranged at a portion of the leading edge of the first wing member 120a near a distal end thereof for better flight control, and the first flap member 128a may be arranged at a portion of the trailing edge of the first wing member 120a so as to be directly opposite (or substantially directly opposite) the first thrust unit 124a (at least substantially under the direct airflow of the first thrust unit 124a) to increase the effectiveness of the first flap member 128a in diverting / controlling the airflow and generating the desired motion in the hybrid aerial vehicle 100. In various embodiments, the second thrust unit 124b and the second flap member 128b are arranged at a leading edge and a trailing edge of the second wing member 120b, respectively. Similarly, the second thrust unit 124b may be arranged at a portion of the leading edge of the second wing member 120b near a distal end thereof, and the second flap member 128b may be arranged at a portion of the trailing edge of the second wing member 120b so as to be directly opposite (or substantially directly opposite) the second thrust unit 124b (at least substantially under the direct airflow of the second thrust unit 124b).

[0018] In various embodiments, the hybrid aerial vehicle 100 further comprises: a frame member arranged along the lateral axis 140 of the central member 110 (e.g., arranged to go through the central member 110). In this regard, the central member 110 is coupled to the frame member (e.g., at a central portion of the frame member, e.g., rotatably coupled thereabout), and the first and second wing members 120a, 120b are rotatably coupled to the frame member such that the first and second wing members 120a, 120b are rotatable about the frame member when actuated by the wing actuator 130.

[0019] In various embodiments, the frame member is a rod, such as a carbon fiber rod.

[0020] In various embodiments, the central member 110 is configured to house the wing actuator 130 and the flight controller 160. For example, the wing actuator 130 and the flight controller 160 may be housed or disposed within or on the central member 110.

[0021] In various embodiments, the flight controller 160 comprises a wing controller configured to: control, in response to the transition control signal being for transitioning the hybrid aerial vehicle 100 from the rotating rotorcraft flight mode to the non-rotating rotorcraft flight mode, the wing actuator 130 to rotate the first and second wing members 120a, 120b about the lateral axis 140 of the central member 110 from an extended state (e.g., the first and second wing members 120a, 120b are angled or oriented at a predetermined angle away from each other) to a retracted state (e.g., the first and second wing members 120a, 120b are angled or oriented parallel (or substantially parallel) to each other), and control, in response to the transition control signal being for transitioning the hybrid aerial vehicle 100 from the nonrotating rotorcraft flight mode to the rotating rotorcraft flight mode, the wing actuator 130 to rotate the first and second wing members 120a, 120b about the lateral axis 140 of the central member 110 from the retracted state to the extended state.

[0022] In various embodiments, the flight controller 160 further comprises: a first flight mode controller configured to output first control output commands for controlling a flight of the hybrid aerial vehicle 100 for the rotating rotorcraft flight mode; a second flight mode controller configured to output second control output commands for controlling a flight of the hybrid aerial vehicle 100 for the non-rotating rotorcraft flight mode; and a mode transition controller communicatively coupled to the first and second flight mode controllers and configured to, for a transition between the rotating rotorcraft flight mode and the non-rotating rotorcraft flight mode, control an amount of the first control output commands and an amount of the second control output commands for contributing in resultant control output commands (e.g., for forming or making up final control output commands) for controlling the flight of the hybrid aerial vehicle 100 based a state of the transition between the rotating rotorcraft flight mode and the non-rotating rotorcraft flight mode.

[0023] In various embodiments, for the transition between the rotating rotorcraft flight mode and the non-rotating rotorcraft flight mode, the amount of the first control output commands and the amount of the second control output commands for contributing in the resultant control output commands are determined based on a wing angle parameter. In this regard, the wing angle parameter corresponds to a collective angle of the first and second wing members 120a, 120b and represents the state of the transition between the rotating rotorcraft flight mode and the non-rotating rotorcraft flight mode.

[0024] In various embodiments, for a transition from the rotating rotorcraft flight mode to the non-rotating rotorcraft flight mode, the amount of the first control output commands andthe amount of the second control output commands for contributing in the resultant control output commands are determined based on a first sigmoid function configured as a function of the wing angle parameter. In various embodiments, for a transition from the non-rotating rotorcraft flight mode to the rotating rotorcraft flight mode, the amount of the first control output commands and the amount of the second control output commands for contributing in the resultant control output commands are determined based on a second sigmoid function configured as a function of the wing angle parameter.

[0025] In various embodiments, the first sigmoid function is parameterized based on one or more first function parameters comprising one or more of a transition midpoint parameter and a transition width parameter for the first sigmoid function. In various embodiments, similarly, the second sigmoid function is parameterized based on one or more second function parameters comprising one or more of a transition midpoint parameter and a transition width parameter for the second sigmoid function.

[0026] In various embodiments, the one or more first function parameters and the one or more second function parameters are optimized based on an objective function.

[0027] In various embodiments, the second flight mode controller is configured to output the second control output commands further for controlling a flight of the hybrid aerial vehicle 100 for the fixed-wing flight mode. That is, the second flight mode controller is configured to output the second control output commands for controlling the flight of the hybrid aerial vehicle 100 for the non-rotating rotorcraft flight mode and for the fixed-wing flight mode. In this regard, the mode transition controller is further configured to, for a transition between the non-rotating rotorcraft flight mode and the fixed-wing flight mode, control an amount of target base pitch angle modulation based a state of the transition between the non-rotating rotorcraft flight mode and the fixed-wing flight mode. For example, the second flight mode controller may be configured to generate the second control output commands for controlling the flight of the hybrid aerial vehicle 100 for the non-rotating rotorcraft flight mode and for the fixed-wing flight mode based on a desired quaternion attitude of the hybrid aerial vehicle 100, and the above- mentioned amount of target base pitch angle modulation modifies the target base pitch angle associated with or provided by the desired quaternion attitude by the above-mentioned amount, prior to being used by the second flight mode controller to generate the second control output commands. Therefore, the second control output commands generated by the second flight mode controller for the non-rotating rotorcraft flight mode and for the fixed-wing flight modeis based on the desired quaternion attitude of the hybrid aerial vehicle 100 as modified by the above-mentioned amount of target base pitch angle modulation.

[0028] In various embodiments, for a transition from the non-rotating rotorcraft flight mode to the fixed-wing flight mode, the amount of target base pitch angle modulation is determined based on a third sigmoid function configured as a function of the state of the transition between the non-rotating rotorcraft flight mode and the fixed-wing flight mode. In various embodiments, for a transition from the fixed-wing flight mode to the non-rotating rotorcraft flight mode, the amount of target base pitch angle modulation is determined based on a fourth sigmoid function configured as a function of the state of the transition between the non-rotating rotorcraft flight mode and the fixed-wing flight mode.

[0029] In various embodiments, the third sigmoid function is parameterized based on one or more third function parameters comprising one or more of a transition midpoint parameter and a transition width parameter for the third sigmoid function. In various embodiments, the fourth sigmoid function is parameterized based on one or more fourth function parameters comprising one or more of a transition midpoint parameter and a transition width parameter for the fourth sigmoid function.

[0030] In various embodiments, the one or more first function parameters, the one or more second function parameters, the one or more third function parameters and the one or more fourth function parameters are optimized (e.g., collectively optimized) based on the objective function.

[0031] In various embodiments, the first control output commands comprise a first flap actuator output command for controlling the first flap actuator 126a, a second flap actuator output command for controlling the second flap actuator 126b, a first thrust output command for controlling the first thrust unit 126a and a second thrust output command for controlling the second thrust unit 126b. In various embodiments, the first flight mode controller is configured to generate the first and second flap actuator output commands of the first control output commands for controlling the flight of the hybrid aerial vehicle 100 for the rotating rotorcraft flight mode based on a pitch angle and a roll angle of a desired tip path plane of the hybrid aerial vehicle 100.

[0032] In various embodiments, similarly, the second control output commands comprise a first flap actuator output command for controlling the first flap actuator 126a, a second flap actuator output command for controlling the second flap actuator 126b, a first thrust output command for controlling the first thrust unit 124a and a second thrust output command forcontrolling the second thrust unit 124b. In various embodiments, the second flight mode controller is configured to generate the second control output commands for controlling the flight of the hybrid aerial vehicle 100 for the non-rotating rotorcraft flight mode and for the fixed-wing flight mode based on a desired quaternion attitude of the hybrid aerial vehicle 100.

[0033] In various embodiments, the second flight mode controller is configured to: generate virtual torque commands based on the desired quaternion attitude of the hybrid aerial vehicle 100; and generate the second control output commands based on the virtual torque commands and a control allocation matrix. In various embodiments, the virtual torque commands may be combined with virtual thrust commands to form combined virtual thrust / torque commands and the control allocation matrix configured to map the combined virtual thrust / torque commands to the second control output commands.

[0034] In various embodiments, the first and second thrust units 124a, 124b may each comprise a propeller and a motor configured to rotate the propeller. It will be appreciated by a person skilled in the art that the present invention is not limited to any particular type of thrust unit for the first and second thrust units 124a, 124b as long as the thrust unit is capable of being operated to generate thrust to facilitate or support a flight of the hybrid aerial vehicle 100 as described herein according to various embodiments. It will also be appreciated by a person skilled in the art that the present invention is not limited to any particular type of actuator for the first and second flap actuators 126a, 126b as long as the flap actuator is capable of being configured, or being operated, to actuate the flap member (e.g., to adjust the flap angle / defl ection) as described herein according to various embodiments. Similarly, the present invention is also not limited to any particular type of linear actuator for the wing actuator 130 as long as the linear actuator is capable of being configured, or being operated, to rotate the first and second wing members 120a, 120b via the first and second mechanical linkages 150a, 150b as described herein according to various example embodiments. Furthermore, it will be appreciated by a person skilled in the art that the wing members 120a, 120b and the flap members 128a, 128b may each be configured to have any configuration or any shape as desired or as appropriate, as long as the hybrid aerial vehicle 100 can operate in the multiple flight modes as described herein according to various embodiments. Accordingly, the wing members 120a, 120b (or wing platform shape) and the flap members 128a, 128b are not limited to any particular configuration or shape. For example, the wing platform shape and / or the placement of the first and second thrust units 124a, 124b and the flap members 128a, 128b may bedetermined via an optimization process, such as to enhance or optimize performance in one or more flight modes.

[0035] FIG. 2 depicts a schematic flow diagram of a method 200 of operating the hybrid aerial vehicle 100 according to various embodiments of the present invention. As described hereinbefore according to various embodiments, the flight controller 160 comprises a wing controller configured to: control, in response to the transition control signal being for transitioning the hybrid aerial vehicle 100 from the rotating rotorcraft flight mode to the nonrotating rotorcraft flight mode, the wing actuator 130 to rotate the first and second wing members 120a, 120b about the lateral axis 140 of the central member 110 from an extended state to a retracted state, and control, in response to the transition control signal being for transitioning the hybrid aerial vehicle 100 from the non-rotating rotorcraft flight mode to the rotating rotorcraft flight mode, the wing actuator 130 to rotate the first and second wing members 120a, 120b about the lateral axis 140 of the central member 110 from the retracted state to the extended state. In this regard, the method 200 comprises: controlling (at 206), by the wing controller and in response to the transition control signal being for transitioning the hybrid aerial vehicle 100 from the rotating rotorcraft flight mode to the non-rotating rotorcraft flight mode, the wing actuator 130 to rotate the first and second wing members 120a, 120b about the lateral axis 140 of the central member 110 from the extended state to the retracted state, and controlling (at 208), by the wing controller and in response to the transition control signal being for transitioning the hybrid aerial vehicle 100 from the non-rotating rotorcraft flight mode to the rotating rotorcraft flight mode, the wing actuator 130 to rotate the first and second wing members 120a, 120b about the lateral axis 140 of the central member 130 from the retracted state to the extended state.

[0036] As described hereinbefore according to various embodiments, the flight controller 160 further comprises: a first flight mode controller configured to output first control output commands for controlling a flight of the hybrid aerial vehicle 100 for the rotating rotorcraft flight mode; a second flight mode controller configured to output second control output commands for controlling a flight of the hybrid aerial vehicle 100 for the non-rotating rotorcraft flight mode; and a mode transition controller communicatively coupled to the first and second flight mode controllers and configured to, for a transition between the rotating rotorcraft flight mode and the non-rotating rotorcraft flight mode, control an amount of the first control output commands and an amount of the second control output commands for contributing in resultant control output commands (e.g., for forming or making up the final control output commands)for controlling the flight of the hybrid aerial vehicle 100 based a state of the transition between the rotating rotorcraft flight mode and the non-rotating rotorcraft flight mode. In this regard, the method 200 further comprises: outputting, by the first flight mode controller, the first control output commands for controlling the flight of the hybrid aerial vehicle 100 for the rotating rotorcraft flight mode; outputting, by the second flight mode controller, the second control output commands for controlling the flight of the hybrid aerial vehicle 100 for the non-rotating rotorcraft flight mode; and controlling, by the mode transition controller and for the transition between the rotating rotorcraft flight mode and the non-rotating rotorcraft flight mode, the amount of the first control output commands and the amount of the second control output commands for contributing in the resultant control output commands for controlling the flight of the hybrid aerial vehicle 100, based the state of the transition between the rotating rotorcraft flight mode and the non-rotating rotorcraft flight mode.

[0037] In various embodiments, the above-mentioned controlling, by the mode transition controller and for the transition between the rotating rotorcraft flight mode and the non-rotating rotorcraft flight mode, the amount of the first control output commands and the amount of the second control output commands for contributing in the resultant control output commands comprises: determining, for the transition between the rotating rotorcraft flight mode and the non-rotating rotorcraft flight mode, the amount of the first control output commands and the amount of the second control output commands for contributing in the resultant control output commands based on a wing angle parameter. In this regard, the wing angle parameter corresponds to a collective angle of the first and second wing members 120a, 120b and represents the state of the transition between the rotating rotorcraft flight mode and the nonrotating rotorcraft flight mode.

[0038] In various embodiments, the above-mentioned determining, for the transition between the rotating rotorcraft flight mode and the non-rotating rotorcraft flight mode, the amount of the first control output commands and the amount of the second control output commands for contributing in the resultant control output commands comprises: determining, for a transition from the rotating rotorcraft flight mode to the non-rotating rotorcraft flight mode, the amount of the first control output commands and the amount of the second control output commands for contributing in the resultant control output commands based on a first sigmoid function configured as a function of the wing angle parameter; and determining, for a transition from the non-rotating rotorcraft flight mode to the rotating rotorcraft flight mode, the amount of the first control output commands and the amount of the second control outputcommands for contributing in the resultant control output commands based on a second sigmoid function configured as a function of the wing angle parameter.

[0039] In various embodiments, the first sigmoid function is parameterized based on one or more first function parameters comprising one or more of a transition midpoint parameter and a transition width parameter for the first sigmoid function. In various embodiments, the second sigmoid function is parameterized based on one or more second function parameters comprising one or more of a transition midpoint parameter and a transition width parameter for the second sigmoid function.

[0040] In various embodiments, the one or more first function parameters and the one or more second function parameters are optimized based on an objective function.

[0041] In various embodiments, the method 200 further comprises: outputting, by the second flight mode controller, the second control output commands for controlling a flight of the hybrid aerial vehicle 100 for the fixed-wing flight mode, and controlling, by the mode transition controller and for a transition between the non-rotating rotorcraft flight mode and the fixed-wing flight mode, an amount of target base pitch angle modulation based a state of the transition between the non-rotating rotorcraft flight mode and the fixed-wing flight mode.

[0042] In various embodiments, the above-mentioned controlling, by the mode transition controller and for the transition between the non-rotating rotorcraft flight mode and the fixed- wing flight mode, the amount of target base pitch angle modulation comprises: determining, for a transition from the non-rotating rotorcraft flight mode to the fixed-wing flight mode, the amount of target base pitch angle modulation based on a third sigmoid function configured as a function of the state of the transition between the non-rotating rotorcraft flight mode and the fixed-wing flight mode; and determining, for a transition from the fixed-wing flight mode to the non-rotating rotorcraft flight mode, the amount of target base pitch angle modulation based on a fourth sigmoid function configured as a function of the state of the transition between the non-rotating rotorcraft flight mode and the fixed-wing flight mode.

[0043] In various embodiments, the third sigmoid function is parameterized based on one or more third function parameters comprising one or more of a transition midpoint parameter and a transition width parameter for the third sigmoid function. In various embodiments, the fourth sigmoid function is parameterized based on one or more fourth function parameters comprising one or more of a transition midpoint parameter and a transition width parameter for the fourth sigmoid function.

[0044] In various embodiments, the one or more first function parameters, the one or more second function parameters, the one or more third function parameters and the one or more fourth function parameters are optimized (e.g., collectively optimized) based on the objective function.

[0045] In various embodiments, the method 200 further comprises: generating, by the first flight mode controller, the first and second flap actuator output commands of the first control output commands for controlling the flight of the hybrid aerial vehicle 100 for the rotating rotorcraft flight mode is based on a pitch angle and a roll angle of a desired tip path plane of the hybrid aerial vehicle 100.

[0046] In various embodiments, the method 200 further comprises generating, by the second flight mode controller, the second control output commands for controlling the flight of the hybrid aerial vehicle 100 for the non-rotating rotorcraft flight mode and the fixed-wing flight mode based on a desired quaternion attitude of the hybrid aerial vehicle 100.

[0047] In various embodiments, the method 200 further comprises: generating, by the second flight controller, virtual torque commands based on the desired quaternion attitude of the hybrid aerial vehicle 100. In various embodiments, the above-mentioned generating, by the second flight controller, the second control output commands are based on the virtual torque commands and a control allocation matrix. In various embodiments, the virtual torque commands may be combined with virtual thrust commands to form combined virtual thrust / torque commands the control allocation matrix configured to map the combined virtual thrust / torque commands to the second control output commands.

[0048] It will be appreciated by a person skilled in the art that the method 200 of operating the hybrid aerial vehicle 100 may further comprise additional or other operation(s) corresponding to operation(s) in which the hybrid aerial vehicle 100 (e.g., components / modules thereof) is configured to perform as described herein according to various embodiments of the present invention, and thus such corresponding additional or other operations need not be repeated with respect to the method 200 of operating the hybrid aerial vehicle 100 for clarity and conciseness. In other words, various embodiments described herein in context of the hybrid aerial vehicle 100 are analogously valid for the method 200 of operating the hybrid aerial vehicle 100, and vice versa.

[0049] FIG. 3 depicts a schematic block diagram of the flight controller 160 configured to control, in response to a transition control signal received, the hybrid aerial vehicle 100 to transition from one flight mode to another flight mode amongst the multiple flight modes duringflight according to various embodiments of the present invention. The flight controller 160 comprises: at least one memory 302; and at least one processor 304 communicatively coupled to the at least one memory 302 and configured to perform the above-mentioned method 200 of operating the hybrid aerial vehicle 100 as described herein according with reference to FIG. 2 according to various embodiments of the present invention.

[0050] It will be appreciated by a person skilled in the art that the at least one processor 304 may be configured to perform various functions or operations through set(s) of instructions (e.g., software modules) executable by the at least one processor 304 to perform various functions or operations. Accordingly, as shown in FIG. 3, the flight controller 160 may comprise: the above-mentioned wing controller 310 (or a wing control module or circuit) configured to: control, in response to the transition control signal being for transitioning the hybrid aerial vehicle 100 from the rotating rotorcraft flight mode to the non-rotating rotorcraft flight mode, the wing actuator 130 to rotate the first and second wing members 120a, 120b about the lateral axis 140 of the central member 110 from the extended state to the retracted state, and control, in response to the transition control signal being for transitioning the hybrid aerial vehicle 100 from the non-rotating rotorcraft flight mode to the rotating rotorcraft flight mode, the wing actuator 130 to rotate the first and second wing members 120a, 120b about the lateral axis 140 of the central member 110 from the retracted state to the extended state. The flight controller 160 may further comprise: the above-mentioned first flight mode controller 312 (or a first flight mode control module or circuit) configured to output first control output commands for controlling a flight of the hybrid aerial vehicle 100 for the rotating rotorcraft flight mode; a second flight mode controller 314 (or a second flight mode control module or circuit) configured to output second control output commands for controlling a flight of the hybrid aerial vehicle 100 for the non-rotating rotorcraft flight mode; and a mode transition controller 316 (or a mode transition control module or circuit) communicatively coupled to the first and second flight mode controllers 312, 314 and configured to, for a transition between the rotating rotorcraft flight mode and the non-rotating rotorcraft flight mode, control an amount of the first control output commands and an amount of the second control output commands for contributing in resultant control output commands (e.g., for forming or making up the final control output commands) for controlling the flight of the hybrid aerial vehicle 100 based a state of the transition between the rotating rotorcraft flight mode and the non-rotating rotorcraft flight mode. It will be appreciated by a person skilled in the art that various components of the flight controller 160 may communicate via an interconnected bus 305.

[0051] It will be appreciated by a person skilled in the art that the wing controller 310, the first flight mode controller 312, the second flight mode controller 314 and the mode transition controller 316 are not necessarily separate modules, and they may be realized by or implemented as one functional module (e.g., a circuit or a software program) as desired or as appropriate without deviating from the scope of the present invention. For example, the wing controller 310, the first flight mode controller 312, the second flight mode controller 314 and / or the mode transition controller 316 may be realized (e.g., compiled together) as one executable software program (e.g., software application or simply referred to as an “app”), which for example may be stored in the at least one memory 302 and executable by the at least one processor 304 to perform various functions / operations as described herein according to various embodiments of the present invention.

[0052] In various embodiments, the flight controller 160 corresponds to the method 200 of operating the hybrid aerial vehicle 100 as described hereinbefore with reference to FIG. 2 according to various embodiments, therefore, various operations, functions or steps configured to be performed by the at least one processor 304 may correspond to various operations, functions or steps of the method 200 of operating the hybrid aerial vehicle 100 as described herein according to various embodiments, and thus need not be repeated with respect to the flight controller 160 for clarity and conciseness. In other words, various embodiments described herein in context of the methods are analogously valid for the corresponding systems (e.g., the flight controller 160), and vice versa.

[0053] For example, in various embodiments, the at least one memory 302 may have stored therein the wing controller 310, the first flight mode controller 312, the second flight mode controller 314 and / or the mode transition controller 316, which are configured to perform, or correspond to, various operations, functions or steps of the method 200 of operating the hybrid aerial vehicle 100 as described herein according to various embodiments, which are executable by the at least one processor 304 to perform the corresponding operations, functions or steps as described herein.

[0054] Any computing system or device providing a processing capability may be provided according to various embodiments of the present invention. Such a computing system or device may be taken to include one or more processors and one or more computer-readable storage mediums. A memory or computer-readable storage medium used in various embodiments may be a volatile memory, for example a DRAM (Dynamic Random Access Memory) or a nonvolatile memory, for example a PROM (Programmable Read Only Memory), an EPROM(Erasable PROM), EEPROM (Electrically Erasable PROM), or a flash memory, e.g., a floating gate memory, a charge trapping memory, an MRAM (Magnetoresistive Random Access Memory) or a PCRAM (Phase Change Random Access Memory).

[0055] In various embodiments, a “circuit” may be understood as any kind of a logic implementing entity, which may be special purpose circuitry or a processor executing software stored in a memory, firmware, or any combination thereof. Thus, in an embodiment, a “circuit” may be a hard-wired logic circuit or a programmable logic circuit such as a programmable processor, e.g., a microprocessor (e.g., a Complex Instruction Set Computer (CISC) processor or a Reduced Instruction Set Computer (RISC) processor). A “circuit” may also be a processor executing software, e.g., any kind of computer program, e.g., a computer program using a virtual machine code, e.g., Java. Any other kind of implementation of various functions or operations may also be understood as a “circuit” in accordance with various other embodiments. Similarly, a “module” may be a portion of a system according to various embodiments of the present invention and may encompass a “circuit” as above, or may be understood to be any kind of a logic-implementing entity therefrom.

[0056] Some portions of the present disclosure may be explicitly or implicitly presented in terms of algorithms and functional or symbolic representations of operations on data within a computer memory. These algorithmic descriptions and functional or symbolic representations are the means used by those skilled in the data processing arts to convey most effectively the substance of their work to others skilled in the art. An algorithm is here, and generally, conceived to be a self-consistent sequence of steps leading to a desired result. The steps are those requiring physical manipulations of physical quantities, such as electrical, magnetic or optical signals capable of being stored, transferred, combined, compared, and otherwise manipulated.

[0057] It will be appreciated by a person skilled in the art that the flight controller 160 may be specially constructed for the required purposes, or may comprise a general purpose computing device selectively activated or reconfigured by computer program(s) stored in the computing device. In general, various algorithms that may be presented herein are not limited to being implemented or executed by any particular computer system. Alternatively, the construction of more specialized computer system to perform various operations, functions or steps of various methods described herein may be provided as desired or as appropriate without going beyond the scope of the present invention.

[0058] In addition, the present specification also at least implicitly discloses computer program(s) or software / functional module(s), in that it would be apparent to a person skilled in the art that various operations, functions or steps of various methods described herein may be put into effect by computer code. The computer program(s) is not intended to be limited to any particular programming language and implementation thereof, and it will be appreciated by a person skilled in the art that a variety of programming languages and coding thereof may be used to implement the computer program(s). Moreover, the computer program(s) is not intended to be limited to any particular control flow as there are a variety of programming languages which can use different control flows. It will be appreciated by a person skilled in the art that a computer program may be stored on any computer-readable storage medium (non- transitory computer-readable storage medium), such as but not limited to, a magnetic disk, an optical disk or a memory chip. For example, a computer program stored on a computer-readable storage medium may be loaded and executed on a computer system or device to implement various operations, functions or steps of various methods described herein according to various embodiments of the present invention.

[0059] Accordingly, in various embodiments, there is provided a computer program product, embodied in one or more computer-readable storage mediums (non-transitory computer-readable storage medium(s)), comprising instructions (e.g., the wing controller 310, the first flight mode controller 312, the second flight mode controller 314 and / or the mode transition controller 316) executable by one or more computer processors to perform the method 200 of operating the hybrid aerial vehicle 100, as described herein with reference to FIG. 2 according to various embodiments of the present invention. Accordingly, various computer programs or software modules described herein may be stored in a computer program product receivable by a system or device therein, such as the flight controller 160, for execution by at least one processor 304 of the flight controller 160 to perform various functions or operations described herein according to various embodiments of the present invention.

[0060] It will be appreciated by a person skilled in the art that various modules described herein (e.g., the wing controller 310, the first flight mode controller 312, the second flight mode controller 314 and / or the mode transition controller 316) may be software module(s) realized by computer program(s) or set(s) of instructions executable by a computer processor to perform various functions or operations. Various modules described herein (e.g., the wing controller 310, the first flight mode controller 312, the second flight mode controller 314 and / or the mode transition controller 316) may also be implemented as hardware module(s)being functional hardware unit(s) designed to perform various functions or operations. More particularly, in the hardware sense, a module is a functional hardware unit designed for use with other components or modules. For example, a module may be implemented using discrete electronic components, or it can form a portion of an entire electronic circuit such as an Application Specific Integrated Circuit (ASIC). Numerous other possibilities exist. It will also be appreciated by a person skilled in the art that a combination of hardware and software modules may be implemented. Furthermore, various operations, functions or steps of various methods described herein may be performed in parallel rather than sequentially as desired or as appropriate (e.g., as long as it does not render the method(s) inoperable or unsatisfactory for its intended purpose).

[0061] FIG. 4 depicts a schematic flow diagram of a method 400 of forming the hybrid aerial vehicle 100 with multiple flight modes as described hereinbefore according to various embodiments the present invention. The method 400 comprises: providing or forming (at 402) a central member 110; providing or forming (at 404) a first wing member 120a arranged at a first side of the central member 110, the first wing member 120a having a first thrust unit 124a and a first flap actuator 126a affixed thereto, the first flap actuator 126a having a first flap member 128a affixed thereto and is configured to actuate the first flap member 128a; providing or forming (at 406) a second wing member 120b arranged at a second side of the central member 110, the second wing member 120b having a second thrust unit 124b and a second flap actuator 126b affixed thereto, the second flap actuator 126b having a second flap member 128b affixed thereto and is configured to actuate the second flap member 128b, whereby the first and second sides of the central member are opposite sides of the central member 110; providing or forming (at 408) a wing actuator 130 configured as a linear actuator and configured to rotate the first and second wing members 120a, 120b about a lateral axis 140 of the central member 110 via a first mechanical linkage 150a and a second mechanical linkage 150b, respectively; and providing or forming (at 410) a flight controller 160 communicatively coupled to the wing actuator 130, the first and second thrust units 124a, 124b and the first and second flap actuators 126a, 126b and configured to control, in response to a transition control signal received, the hybrid aerial vehicle 100 to transition from one flight mode to another flight mode amongst the multiple flight modes during flight based on controlling the wing actuator 130, the first and second thrust units 124a, 124b and the first and second flap actuators 126a, 126b. In particular, the multiple flight modes comprise a rotating rotorcraft flight mode, a non-rotating rotorcraft flight mode and a fixed-wing flight mode.

[0062] The method 400 is for forming the hybrid aerial vehicle 100 as described hereinbefore with reference to FIG. 1, therefore, various steps or operations of the method 400 may correspond to forming, providing or configuring various components, modules or portions of the hybrid aerial vehicle 100 as described herein according to various embodiments, and thus such corresponding steps or operations need not be described or repeated with respect to the method 400 for clarity and conciseness. In other words, various embodiments described herein in context of the hybrid aerial vehicle 100 are analogously valid for the method 400 of forming the hybrid aerial vehicle 100 having various components, modules, portions and configurations as described hereinbefore according to various embodiments, and vice versa. It will also be appreciated by a person skilled in the art that the method 400 of forming the hybrid aerial vehicle 100 is not limited to any particular order of operations / steps. For example, FIG. 4 does not indicate or limit any particular order of operations / steps in which the method 400 can be performed to form the hybrid aerial vehicle 100. Furthermore, one or more operations / steps of the method 400 may be performed concurrently or integrally as desired or as appropriate without going beyond the scope of the present invention. For example, various components or parts of the hybrid aerial vehicle 100 may be 3D printed components or parts, such as but not limited to, the central member 110, the first and second wing members 120a, 120b, the first and second flap member 128a, 128b, and may be assembled together (e.g., coupled together), along with various other components or parts of the hybrid aerial vehicle 100, to form the hybrid aerial vehicle 100 as described herein according to various embodiments of the present invention.

[0063] It will be appreciated by a person skilled in the art that the terminology used herein is for the purpose of describing various embodiments only and is not intended to be limiting of the present invention. As used herein, the singular forms “a”, “an” and “the” are intended to include the plural forms as well, unless the context clearly indicates otherwise. It will be further understood that the terms “comprises” and / or “comprising,” when used in this specification, specify the presence of stated features, integers, steps, operations, elements, and / or components, but do not preclude the presence or addition of one or more other features, integers, steps, operations, elements, components, and / or groups thereof.

[0064] Any reference to an element or a feature herein using a designation such as “first”, “second” and so forth does not limit the quantity or order of such elements or features, unless stated or the context requires otherwise. For example, such designations may be used herein as a convenient way of distinguishing between two or more elements or instances of an element.Thus, a reference to first and second elements does not necessarily mean that only two elements can be employed, or that the first element must precede the second element, unless stated or the context requires otherwise. In addition, a phrase referring to “at least one of’ a list of items refers to any single item therein or any combination of two or more items therein.

[0065] In order that the present invention may be readily understood and put into practical effect, various example embodiments of the present invention will be described hereinafter by way of examples only and not limitations. It will be appreciated by a person skilled in the art that the present invention may, however, be embodied in various different forms or configurations and should not be construed as limited to the example embodiments set forth hereinafter. Rather, these example embodiments are provided so that this disclosure will be thorough and complete, and will fully convey the scope of the present invention to those skilled in the art.

[0066] Within the realm of rotorcrafts, there exists a type of rotor-wing craft which spins its entire body around a central axis of rotation to generate lift, essentially turning the craft into one big propeller, which may be referred to as a rotating rotorcraft. The single-winged version, called a monocopter, draws its inspiration from the falling of the samara seed, which gracefully rotates to the ground in a process called autorotation. For example, FIG. 5 A shows example maple seeds and FIG. 5B show a composite image of the autorotation motion of the maple seed. By extending this monocopter rotor-wing approach to two symmetrical wings, various example embodiments fuse the rotating rotor-wing approach (e.g., corresponding to the “rotating rotorcraft flight mode” as described hereinbefore according to various embodiments) with a fixed-wing approach (e.g., corresponding to the “fixed-wing flight mode” as described hereinbefore according to various embodiments, e.g., a tailless fixed-wing approach) and a dualcopter approach whereby the wings are oriented vertically / perpendicularly (or substantially vertically / perpendicularly) to the ground during hover (e.g., corresponding to the “non-rotating rotorcraft flight mode” as described hereinbefore according to various embodiments) to form a versatile and agile multi-modal hybrid craft, and thus with improved operational efficiency and effectiveness. FIG. 5C depicts a multi-modal hybrid aerial vehicle under the rotating rotorcraft flight mode according to various example embodiments of the present invention to illustrate the rotating motion which generates lift via spinning about a central axis.

[0067] Various example embodiments note that one of the challenges of building a hybrid system is the additional requirement to have components that would work in all flight regimes, for example having a propulsion system that works well both in a rotorcraft system (e.g.,corresponding to the “rotating rotorcraft flight mode” and the “non-rotating rotorcraft flight mode”) and a forward flight system (e.g., corresponding to the “fixed-wing flight mode”). A way to circumvent this is to have redundant components at the expense of additional weight, for example, separate propulsion systems. However, various example embodiments note that it is advantageous to have structurally efficient designs, whereby sub-components are utilized in all flight modes / regimes, for example, using the same wings for hovering and cruising. This structural efficiency can be realized when combining the rotating rotorcraft approach and the fixed-wing approach, as the rotating rotorcraft approach uses the rotating wings for lift instead of directly using the propulsion system unlike a rotorcraft (e.g., a multirotor system) operating under the non-rotating approach, where the wings would not be used to generate lift and constitute redundant weight. In various example embodiments, these same wings used during the rotating rotorcraft flight mode are also used during the fixed-wing flight mode, thus maximizing their use.

[0068] Even though the long-distance capability of a non-rotating rotorcraft (e.g., a multirotor system) may be limited, various example embodiments note its advantage in maneuverability, and pose several advantages over a rotating rotorcraft, such as having a nonrotating frame and also increased agility to navigate tight spaces. While the efficiency of the rotating hover mode cannot be matched by the non-rotating hover mode of a non-rotating rotorcraft, various example embodiments note that providing the non-rotating rotorcraft flight mode advantageously provides the user with an on-demand option to enter the rotorcraft into a more agile state at the expense of power efficiency (drawing more power). For example, in an example hybrid dual-wing system, various example embodiments realize the non-rotating rotorcraft flight mode as a flight mode whereby the flight of the hybrid dual-wing aerial vehicle (e.g., lift thereof) is primarily achieved by thrust units (e.g., motor-driven propellers) without necessarily rotating / spinning the entire hybrid dual-wing aerial vehicle about its yaw axis, such as in a similar manner as a tailsitter type aircraft. Accordingly, the hybrid dual-wing aerial vehicle uses both the wing flaps and the thrust units for flight control. In various example embodiments, a differential thrust configuration may be employed which increases the agility of the hybrid aerial vehicle while cruising, for example, allowing for smaller turning radii without excessively losing altitude.

[0069] Accordingly, various example embodiments provide a three-mode hybrid aerial vehicle (e.g., UAV) (e.g., corresponding to the hybrid aerial vehicle 100 with multiple flight modes as described hereinbefore according to various embodiments) which possesses thecapability and option to hover efficiently as a rotorcraft under the rotating rotorcraft flight mode or to hover more agilely as a rotorcraft under the non-rotating rotorcraft flight mode (e.g., like a tailsitter whereby the wings are oriented vertically / perpendicularly (or substantially vertically / perpendicularly) to the ground during hover), while also having the ability to transit longer distances efficiently at higher speeds by operating as a fixed-wing under the fixed-wing flight mode. For example, such a three-mode UAV is advantageously employable in a wider range of practical application, such as depending on the mission or task profile, and is adaptable in an operating environment during flight (e.g., adapt quickly while in the air). For illustration purpose only, FIGs. 6A to 6C show an overview of the three flight modes of the hybrid UAV, namely, rotating rotorcraft flight mode, non-rotating rotorcraft flight mode and fixed-wing flight mode, respectively, according to various example embodiments of the present invention. In various example embodiments, in the rotating rotorcraft flight mode (e.g., denoted as the R- mode), the wings are positioned or oriented so as to be angled away from each other as shown in FIG. 6A. In the non-rotating rotorcraft flight mode or the fixed-wing flight mode, the wings are positioned or oriented so as to be parallel (or substantially parallel) to each other. The nonrotating rotorcraft flight mode (e.g., denoted as the T-mode) may also be referred to as a transition mode since the UAV is configured to first transition to the non-rotating rotorcraft flight mode as an intermediate step or phase when transitioning from the rotating rotorcraft flight mode to the fixed-wing flight mode and vice versa. The non-rotating rotorcraft flight mode may also be referred to as a tailsitter mode since it maneuvers like a tailsitter. The fixed- wing flight mode (e.g., denoted as the C-mode) may also be referred to as a cruise mode.

[0070] Various existing multi-modal UAVs will now be discussed. For example, as discussed in the background, one of the simplest forms of hybrid UAV is where a quadrotor frame is simply fused with a fixed-wing frame. Although simple, such a hybrid UAV is not very structurally efficient due to redundant components. For example, one of the earlier works on hybrid UAV is the T-Wing, which proved the feasibility of tail sitter-type aircraft where the control surfaces are submerged in the propeller wash of the propulsion system. They found via optimization that for their platform, the thrust-to-weight ratio used can be as low as 1.15. Some other previous works include a tilt-rotor UAV design that actuates the angle of the propulsion system to accommodate the different flight modes. For example, there has been disclosed a quadrotor tail sitter for autonomous parcel delivery. This quadrotor tail sitter has four motors with no wing control surfaces, and is controlled purely by differential thrust. For example, commercially available hybrid UAV designs include the FireFly6 by BirdsEyeView Aeroboticsand WingtraOne by Wingtra. However, these hybrid aerial vehicles with two flight modes mostly draw their efficiency from the fixed-wing flight mode, and only use the hover mode as a means for vertical take-off and landing. In contrast, various example embodiments enable at least two efficiency regions (at least efficient regions in hovering and cruising) and yet do not compromise on maneuverability by additionally providing a third flight mode, namely, the nonrotating rotorcraft flight mode (the T-mode) for hovering (e.g., like a tailsitter whereby the wings are oriented vertically / perpendicularly (or substantially vertically / perpendicularly) to the ground during hover).

[0071] There has been disclosed a number of rotor-wing hybrid platforms as well as various challenges associated with such platforms, which have resulted in the cancellation of prototype developments, for example, the Boeing X-50 which relied on a complicated tip-jet-driven mechanisxm to the main rotor.

[0072] In various example embodiments, the rotating rotorcraft approach (the R-mode) is utilized to provide three-mode capability including efficient hover. For example, the single winged version is known as a monocopter. Monocopters have been gaining popularity as electronics are miniaturized and more widely available. Earlier works on monocopters have evaluated their feasibility in flight, and demonstrated their flight operation in confined spaces. For example, there has been disclosed a monocopter prototype which could fly across a 70 m by 3 m corridor followed by a 1.5 m doorway. A link between helicopter dynamics and control with this nature-inspired design was also established. However, carrying a payload such as a camera proved to be challenging for a monocopter due to video imagery being sensitive to rotation. Software corrections and solutions have been proposed. Collision avoidance and autonomous capability were also implemented and demonstrated in an indoor environment.

[0073] There has been disclosed a hybrid UAV with two flight modes which utilized two direct high-speed servos in the central hub connected to the wings which was used as both the collective angle and cyclic control mechanism. Various example embodiments note that this dual purpose required the servos to be extremely fast to be able to keep up with the rotation of the UAV, as well as have a large enough rotation range for transition purposes, and at the same time have sufficient torque to withstand the forces during rotation / transition. This limited the choice of servos that fit such criteria. In addition, the transition mechanism was easily damaged and the servo overloaded in the event of a hard landing. In contrast, various example embodiments address such deficiencies or limitations by using a linear actuator comprising an internal worm-gear mechanism for the transition system of the hybrid UAV.

[0074] In another previous work, the basic design principles of the above-mentioned hybrid UAV with two flight modes were examined and a more in-depth theoretical analysis was conducted, as well as explored a transition control strategy based on trajectory optimization to populate pre-computed lookup tables which were referred to by the transition controller. This transition control was only implemented offboard within a controlled environment. A previous study on a joint optimization process for the UAV design and control during its rotary hover mode had also been conducted. In this previous study, the UAV wing design and component placements were optimized and a more modular control framework was devised for the hover mode, where its parameters were also optimized in the same process. However, the transition and cruise capabilities were not fully explored further. The control architecture presented in the previous study also only encompasses the hover mode. In addition, the control was only implemented offboard.

[0075] Accordingly, previous works rely on separate controllers / electronics hardware to demonstrate or realise the different flight modes and also only encompassed two flight modes at most, such as hover and cruise. In contrast, various example embodiments integrate the control architecture for different flight modes, including transition control techniques, into a single hardware platform with onboard control, and extend the operational capability to three flight modes, namely, the rotating rotorcraft flight mode (the R-mode), the non-rotating rotorcraft flight mode (the T-mode) and the fixed-wing flight mode (the C-mode). In this regard, various example embodiments advantageously provide the full control architecture for the entire flight envelope of the hybrid UAV, includinx all three flight modes and their respective transitions.

[0076] FIG. 7 depicts a schematic drawing of a hybrid UAV 700 with three flight modes according to various embodiments of the present invention (e.g., corresponding to the hybrid aerial vehicle 100 as described hereinbefore according to various embodiments). The hybrid UAV 700 comprises: a central member 710; a first wing member 720a (which may simply be referred to herein as a first wing) arranged at a first side (e.g., left side) of the central member 710, the first wing 720a having a first thrust unit 724a and a first flap actuator 726a (which may also be referred to herein as a first flap servo) affixed thereto, the first flap actuator 726a having a first flap member 728a (which may simply be referred to herein as a first flap) affixed thereto and is configured to actuate the first flap 728a (e.g., to adjust the flap angle / deflection); a second wing 720b (which may simply be referred to herein as a first wing) arranged at a second side (e.g., right right) of the central member 710, the second wing 720b having a second thrust unit724b and a second flap actuator 726b (which may also be referred to herein as a second flap servo) affixed thereto, the second flap actuator 726b having a second flap member 728b (which may simply be referred to herein as a second flap) affixed thereto and is configured to actuate the second flap 728b (e.g., to adjust the flap angle / deflection), whereby the first and second sides of the central member are opposite sides (e.g., left and right sides) of the central member 710; a wing actuator (not shown in FIG. 7, e.g., housed in the central member 710) configured as a linear actuator and configured to rotate the first and second wings 720a, 720b about a lateral axis 740 of the central member 710 via a first mechanical linkage 750a and a second mechanical linkage 750b, respectively; and a flight controller (not shown in FIG. 7, e.g., housed on or in the central member 710) communicatively coupled to the wing actuator, the first and second thrust units 724a, 724b and the first and second flaps 726a, 726b and configured to control, in response to a transition control signal received, the hybrid UAV 700 to transition from one flight mode to another flight mode amongst the multiple flight modes during flight based on controlling the wing actuator, the first and second thrust units 724a, 724b and the first and second flaps 726a, 726b. In particular, the multiple flight modes comprise a first flight mode (corresponding to the rotating rotorcraft flight mode), a second flight mode (corresponding to the non-rotating rotorcraft flight mode) and a third flight mode (corresponding to the fixed- wing flight mode).

[0077] In various example embodiments, the first and second thrust units 724a, 724b may each be placed near the distal end of the corresponding wing 720a / 720b for better flight responsiveness, such as in changing the rotation speed of the UAV 700 in the R-mode, and the first and second flaps 728a, 728b may each be placed directly (or substantially directly) under the airflow of the corresponding thrust unit 724a / 724b to increase the effectiveness of the flaps in diverting / controlling the airflow and generating the desired motion in the UAV 700.

[0078] In various example embodiments, the first or rotating rotorcraft flight mode (denoted as the R-mode) is a flight mode whereby the UAV 700 is operating as a rotating rotorcraft and the flight of the UAV 700 (e.g., lift thereof) is primarily achieved by rotating / spinning the entire UAV 700 (with its wing members 720a, 720b (e.g., rotating blades)) about its yaw axis through thrust generated by the thrust units 724a, 724b (e.g., motor-driven propellers). In this regard, control may be achieved by actuating the wing flaps 728a, 728b in an oscillatory manner. The second or non-rotating rotorcraft flight mode (denoted as the T- mode) is a flight mode whereby the UAV 700 is operating as a non-rotating rotorcraft and the flight of the UAV 700 (e.g., lift thereof) is primarily achieved by the thrust units 724a, 724bwithout necessarily rotating / spinning the entire UAV 700 about its yaw axis. For example, in the second flight mode, the UAV 700 may hover or maneuver like a tail sitter and purely uses the thrust units 724a, 724b for lift. In the second flight mode, the attitude of the UAV 700 is more maneuverable than the first flight mode due to the non-rotational nature. This second flight mode may also be used as an intermediary step or mode for transiting from the first flight mode to the third flight mode, or from the third flight mode to the first flight mode. The third or fixed-wing flight mode (denoted as the C-mode) is a flight mode whereby the UAV 700 is operating as a fixed-wing aircraft and the flight of the UAV 700 (e.g., lift thereof) is primarily achieved from the wings 720a, 720b by maintaining a forward velocity through thrust generated by the thrust units 724a, 724b (e.g., as a tailless fixed-wing aircraft).

[0079] For illustration purpose and by way of examples only, FIG. 8 depicts the UAV 700 in each of the above-mentioned three flight modes, and transitions therebetween, according to various example embodiments of the present invention. As shown in FIG. 8, to transit between the R-mode and the T- or C-mode, the collective angle (e.g., denoted by Ω in FIG. 9B) of the wings 720a, 720b is mechanically actuated by a central servo mechanism (i.e., the above- mentioned wing actuator). In the R-mode, the wings 720a, 720b are angled away from each other (e.g., each angled by a predetermined angle Ω from a horizontal / lateral plane of the UAV 700 and such that the leading edge moves towards the direction of rotation of the UAV 700). On the other hand, in the T- and C-mode, the wings 720a, 720b are angled parallel to each other and along a vertical plane (perpendicular to the horizontal / lateral plane) of the UAV 700, for example, enabling a forward flight in the C-mode. Accordingly, in various example embodiments, in response to a transition control signal being for transitioning the UAV 700 from the R-mode to the T-mode, the wing actuator may be controlled to rotate the wings 720a, 720b about the lateral axis 740 of the central member 710 from an extended state to a retracted state, such as shown in FIG. 8, as well as FIG. 9A. Accordingly, in various example embodiments, in response to a transition control signal being for transitioning the UAV 700 from the T-mode to the R-mode, the wing actuator may be controlled to rotate the wings 720a, 720b about the lateral axis 740 of the central member 710 from the retracted state to the extended state.

[0080] In various example embodiments, a transition strategy or technique is provided which involves a transition coordinator that manages the transition sequence, and sigmoid blending weights that control the influence each of the different controllers (e.g., the first flightmode controller 312 / 1112 and the second flight mode controller 314 / 1114) have on the final control output commands for controlling a flight of the UAV 700.

[0081] For better understanding, a dynamic model of the UAV 700 will now be described according to various example embodiments of the present invention. The two main reference frames used are the inertial frame J and the UAV body frame β, made up of orthogonal bases . as shown in FIG. 9A. The body frame is tagged to thecentral hub of the UAV 700.

[0082] According to various example embodiments, the UAV state may be described as a 13 -dimensional vector , representing position in the inertial frame J ,velocity in body frame β , quaternion attitude, and angular velocity in body frame β respectively. In various example embodiments, the UAV 700 has five control inputs which may be described as a vector: , whereby Um1corresponds to athrust power for the first thrust unit 724a (e.g., motor 1) (in the direction of the body frameB), Um2corresponds to a thrust power for the second thrust unit 724b (e.g., motor 2) (in the direction of the body frame B), corresponds to a direction or orientation for the firstflap member 728a (e.g., flap 1) (in the direction of the body frame B), corresponds toa direction or orientation for the second flap member 728b (e.g., flap 2) (in the directionof the body frame B) and Uθcorresponds to a collective angle (e.g., denoted by Ω in FIG. 9B) of the wings 720a, 720b for the wing actuator. In various example embodiments, control inputs Um1, Um2, Uθ(i.e., motors inputs and collective servo input) may be normalized from 0 to 1 (i.e., Um1, Um2, Uθ∈ [0, 1] ), and control inputs (i.e., flap servo inputs) may be normalized from -1 to 1 (i.e., Uf1Uf2∈ [—1, 1]), where 0 indicates a neural position (i.e., parallel to the wings).

[0083] In various example embodiments, the dynamic model used for numerical simulation for the purpose of predicting flight maneuvers and performance is based on 6-DOF equations of motion where the state derivatives can be described as , where:(Equation 1)(Equation 2)

[0084] In various example embodiments, wing aerodynamic forces are modelled using Blade Element Theory (BET). In this regard, the wing is sliced into a number (n) of bladeelements. Individual force and torque components of each element are summed to obtain the total force contribution of the wing. Each blade element’s lift and drag may be defined as:(Equation 3)(Equation 4)

[0085] In the above equation, U∞refers to the incident wind velocity (Ux,eff, Uyare its components), αeffdenotes the effective angle of attack, 8 denotes the flap deflection angle, and dA denotes the area of the blade element. Accordingly, Equation (4) determines the total effective wind velocity acting on the wing, taking into account the effect of the thrust of the thrust unit. Uydenotes the vertical component of the incident wind velocity (e.g., velocity of the blade element coordinate system with respect to the stationary world frame coordinate). Accordingly, the propeller inflow on the wing and its effects has been taken into account in the velocity via the Momentum Theory, and this interaction is captured in Equation (4), where FMdenotes the thrust force of the thrust unit, ρairdenotes the density of air and Apropdenotes the area of the propeller rotor disc. The inflow velocity of each blade element may be individually calculated with respect to the inertial frame, which ensures that the different velocities experienced by each blade element when the UAV 700 is in different states (e.g., hovering, climbing, moving, etc.) are taken into account. If the blade element corresponds to a section of the wing which has no flap, the flap deflection angle 8 may simply be set to zero. Cland Cddenote lift and drag coefficients, respectively, which may be defined as functions of ae^ and 8. As an example illustration, FIG. 9B illustrates aerodynamic forces on a single blade element according to various example embodiments of the present invention.

[0086] Accordingly, in various example embodiments, the force contribution (FWi) and the torque contribution (MWi) of the i-th wing 720a / 720b, where i = { 1, 2}, can be calculated as:(Equation 5) where dFk= dL + dD, k is the &-th blade element of that particular wing and rkrefers to the distance between the 14 chord point (from the leading edge of the wing) of the blade element and the origin of the body frame coordinate system such as shown in FIG. 9B.

[0087] FIG. 10A depicts a cross-sectional front view of the UAV 700 (in the T- or C-mode position), excluding the wings 720a, 720b for visual clarity, according to various example embodiments of the present invention. As described hereinbefore, according to various example embodiments, the UAV 700 comprises a central member 710 (e.g., a central hub or a housing member) connected to two wings 720a, 720b (e.g., two symmetrical wings) on opposite sides (e.g., left and right sides). The central hub 710 may be configured to house both the wing actuator 730 (in the form of a linear actuator) and various flight control electronics, including the flight controller (not shown in FIG. 10 A) described herein. In various example embodiments, the first and second mechanical linkages 750a, 750b (e.g., at respective first ends thereof) are coupled to the first and second wings 720a, 720b, respectively, and are each (e.g., at a respective second end thereof, opposite to the respective first end) further coupled to the wing actuator 730. In various example embodiments, as shown in FIG. 10B, the first and second mechanical linkages 750a, 750b are each configured to transform a linear motion of the wing actuator 730 (e.g., at the respective first end of the first / second mechanical linkage 750a / 750b) to a rotational motion (e.g., at the respective second end of the first / second mechanical linkage 750a / 750b) for rotating the corresponding wing member 720a, 720b about the lateral axis 740 of the central hub 710. In various example embodiments, the wing actuator is configured (together with the first and second mechanical linkages 750, 750b) to rotate the first and second wings 720a, 720b about the lateral axis 740 of the central hub 710 in opposite directions (e.g., clockwise and anti-clockwise directions) via the first and second mechanical linkages 750a, 750b, respectively. In various example embodiments, as can be seen in FIG. 7 for example, the first thrust unit 724a and the first flap 728a are arranged at a leading edge and a trailing edge of the first wing 720a (e.g., proximal to the distal end of the first wing 720a), respectively. In various example embodiments, the second thrust unit 724b and the second flap 728b are arranged at a leading edge and a trailing edge of the second wing 720b (e.g., proximal to the distal end of the second wing 720b), respectively. It will be appreciated by a person skilled in the art that the placement of the first and second thrust units 724a, 724b along the leading edge and the first and second flaps 728a, 728b along the trailing edge may be configured as appropriate, such as optimized to enhance flight performance.

[0088] In various example embodiments, as shown in FIG. 10A, the UAV 700 further comprises: a frame member 760 arranged along the lateral axis 740 of the central hub 710. In this regard, the central hub 710 is coupled to the frame member 760 (e.g., at a central portion of the frame member, e.g., rotatably coupled thereabout), and the first and second wings 720a,720b are rotatably coupled to the frame member 760 such that the first and second wings 720a, 720b are rotatable about the frame member 760 when actuated by the wing actuator. In various example embodiments, the frame member 760 is a rod, such as a carbon fiber rod. Accordingly, a single rod may be arranged to run through the body of the UAV 700 along the lateral axis 740 and connects and provides structural support for the two wings 720a, 720b and the central hub 710 together, as well as acting as a pivot point for the wings 720a, 720b to change angle when actuated. For example, additional fixtures may be provided at the root of each wing 720a / 720b to prevent the wing from unwanted rotation. For example, the additional fixtures may be configured to ensure that the two wings 720a, 720b always move together with the mechanical linkages 750a, 750b and do not freely rotating on their own. In various example embodiments, the flap servos 726a, 726b may be mounted in a 3D-printed housing and directly drive the flaps 728a, 728b. Batteries 764 may also be disposed near the wing root on both sides, such as illustrated in FIG. 10 A.

[0089] Accordingly, in various example embodiments, a single linear actuator 730 is advantageously provided in the central hub 710 of the UAV 700 for controlling the collective angle of both the wings 720a, 720b. Furthermore, linkage mechanisms 750a, 750b are provided to transform the linear motion of the wing actuator 730 to a rotational motion, such as shown in FIG. 10B. In particular, FIG. 10B depicts a schematic drawing of the mechanical linkage 750a / 750b showing the relations between various components thereof when the linear actuator 730 (and thus the wings 720a, 720b) is in an extended state (e.g., the wings 720a, 720b are extended or protruded to form a predetermined angle with a horizontal / lateral plane of the UAV 700) during hover in the R-mode and when the linear actuator 730 (and thus the wings 720a, 720b) is in a retracted state (e.g., the wings are parallel to each other and along a vertical plane (perpendicular to the horizontal / lateral plane) of the UAV 700) during the T- or C-mode. FIG. 10C depicts a schematic drawing of a cross-section of the mechanical linkage 750a / 750b in the extended state. By way of an example only and without limitation, as shown in FIG. 10C, each mechanical linkage 750a / 750b may comprise a first link Li and a second link L2. The wing actuator 730 (configured as a linear actuator) may be supported by and secured to the central hub 710 and comprises a slider joint configured to move in a linear manner (i.e., along one axis) between two ends (e.g., vertically up and down along a vertical axis). One end of the first link Li may be connected / coupled to the slider joint by a revolute joint (thus, freely rotatable in one axis). The other end of the first link Li may be coupled by a revolute joint to one end of the second link L2. The other end of the second link L2 may be coupled to a revolute joint realizedby the rod 760. For example, as the wing actuator 730 moves from the retracted state to the extended state, for each mechanical linkage 750a / 750b, the first link Li pulls the second link L2 and induces a rotation about the revolute joint realized by the rod 760, which in turn rotates the corresponding wing 720a / 720b from the T- or C-mode configuration to the R-mode configuration. In particular, the above-mentioned other end of the second link L2 is physically coupled and mechanically linked to the corresponding wing 720a / 720b about the revolute joint realized by the rod 760, such that a rotation of the second link L2 about the rod 760 would induce the same amount of rotation on the corresponding wing 720a / 720b. On the other hand, when the wing actuator 730 moves from the extended state to the retracted state, the first link Li pushes down on the second link L2, which in turn rotates the corresponding wing 720a / 720b from the R-mode to the T- or C-mode configuration, such that the wings 720a, 720b are parallel to each other.

[0090] By way of an example only and without limitation, a linear actuator (servo) 730 having a stroke length of 30 mm and a gear ratio 210: 1 may be used for controlling the collective angle of the two wings 720a, 720b. Based on such an example linear actuator 730, for example, it may take about 6 seconds for the wings 720a, 720b to move from the fully retracted state to the fully extended state and vice versa. For example, such a linear actuator 730 was found to have minimal backlash and high actuation torque.

[0091] In various example embodiments, an input of the linkage mechanism 750a / 750b is the actuator extension distance / length (D) of the linear actuator 730. In this regard, the actuator extension distance / length corresponds to the distance between the wing pivot point (i.e., the rod 760) to the position of the slider joint (corresponding to an end of the linear actuator 730). Example equations of motion describing the relation of the actuator extension distance D to the output wing collective angle (e.g., denoted by 0 in FIG. 10C) will now be described according to various example embodiments of the present invention. The linkage mechanism 750a / 750b comprises a plurality of links and a plurality of revolute joints, and may be configured akin to a slider-crank mechanism. In various example embodiments, the first and second links Zi, Li may have different lengths, for example, the first linkZi may be longer than the second linkZ2. According to various example embodiments, the position of the point (or revolute joint) linking the first and second links Zi, Z2 together may be described as a system of two equations with two unknowns, as follows:(Equation 6)

[0092] Solving for the actuator extension length D,(Equation 7)

[0093] Thus, minimum and maximum actuator extension length D for a wing angle operating range of Θ ∈ [0, π / 2] may be expressed in terms of the linkage lengths as follows:(Equation 8)

[0094] FIG. 11A depicts a schematic block diagram of the flight controller 1160 of the UAV 700 according to various example embodiments of the present invention. In particular, FIG. 11A shows various data flow as well as an example control architecture of the flight controller 1160. As shown, the raw control signal from an operator / user may be converted to desired states and fed to the R-mode attitude controller 1112 (e.g., corresponding to the first flight mode controller 312 as described hereinbefore according to various embodiments) and the T / C-mode attitude controller 1114 (e.g., corresponding to the second flight mode controller 314 as described hereinbefore according to various embodiments). In various example embodiments, as will be described later below, the R-mode attitude controller 1112 may be configured based on a Tip Path Plane (TPP) of the UAV 700 and a Central Pattern Generator (CPG) 1122, and thus may be referred to as a R-mode TPP / CPG attitude controller. In addition, as shown in FIG. 11 A, the flight controller 1160 comprises a mode transition controller 1116 comprising a control blending block 1117 communicatively coupled to the R-mode attitude controller 1112 and the T / C-mode attitude controller 1114 for mixing or blending the control output commands (e.g., corresponding to the “first control output commands” and the “second control output commands” as described hereinbefore according to various embodiments) therefrom depending on the transition state, which may be based on or linked to the collective angle of the wings 720a, 720b. The mode transition controller 1116 further comprises a C- transition control block 1118 (or simply referred to as a C-control block, which may also be referred to as a base pitch angle modulation block) configured to, for a transition between theT-mode and the C-mode, modify the target base pitch angle associated with or provided by the desired quaternion attitude prior to the desired quaternion attitude (as modified) being used by the T / C-mode attitude controller 1114 to generate the control output commands. In particular, the C-transition control block 1118 may be configured to control an amount of target base pitch angle modulation (by directly modifying the desired quaternion attitude) based on a state of the transition between the T-mode and the C-mode. Therefore, for a transition between the T-mode and the C-mode, the control output commands generated by the T / C-mode attitude controller 1114 for the T-mode and for the C-mode is based on the desired quaternion attitude of the UAV 700 as modified by the above-mentioned amount of target base pitch angle modulation.

[0095] Accordingly, in various example embodiment, the flight controller 1160 comprises a wing controller 1111 configured to control, in response to the transition control signal being for transitioning the UAV 700 from the R-mode to the T-mode, the wing actuator 730 to rotate the wings 720a, 720b about the lateral axis 740 of the central member 710 from an extended state to a retracted state. In addition, the wing controller 1111 is configured to control, in response to the transition control signal being for transitioning the UAV 700 from the T-mode to the R-mode, the wing actuator 730 to rotate the wings 720a, 720b about the lateral axis 740 of the central member 710 from the retracted state to the extended state. The flight controller 1160 may further comprise an altitude controller 1113 configured to control an altitude of the UAV 700 based on a desired altitude and a current altitude of the UAV 700.

[0096] In various example embodiments, the altitude controller 1113 may be configured to generate and output thrust commands (e.g., virtual thrust commands Vthrust) for controlling the altitude of the UAV 700 based on the desired altitude and a current altitude of the UAV 700. For example, the altitude controller 1113 may be configured to generate the thrust commands Vthrust) for the first and second thrust units 724a, 724b as follows: Vthrust=PlD(Altsp, Alt)(Equation 9)

[0097] For example, the altitude controller 1113 comprises a PID controller 1115 configured to receive a current altitude (Alt) and a desired altitude (desired altitude setpoint Altsp) of the UAV 700 and generate thrust commands Vthrustto reach the desired altitude setpoint Altsp. In this regard, as shown in FIG. 11A, the altitude controller 1113 is communicatively coupled to the R-mode attitude controller 1112 and the T / C-mode attitude controller 1114 and is configured to output the thrust commands Vthrustgenerated to the R- mode attitude controller 1112 and the T / C-mode attitude controller 1114.Tip Path Plane / Central Pattern Generator Controller (R-Mode)

[0098] FIG. 1 IB depicts a schematic block diagram of the R-mode attitude controller 1112 according to various example embodiments of the present invention, which shows an example control flow of the R-mode attitude controller 1112 for the UAV 700 in the R-mode configuration. The desired Tip Path Plane (TPP) pitch and roll angles (setpoints sp) (Γpitcft spand Γroll sp) may be inferred from the manual control input (e.g., Radio Control (RC) Controller). For the R-mode, according to various example embodiments, a lower level Central Pattern Generator (CPG)-based controller 1122 may be employed to generate the required oscillatory commands (flap actuator output commands ) for the flap actuators 726a, 726b from a higher level virtual control input. In addition, the R-mode attitude controller 1112 may further comprise a thrust command scaler 1124 configured to scale the virtual thrust commands Vthrust received from the altitude controller 1113 to obtain the output thrust commands (Um1,m2(i .e , um1, um2) for the first and second thrust units 724a, 724b. In various example embodiments, the output thrust commands um1, um2are directly proportional to the thrust commands Vthrustfrom the altitude controller 1113 according to a linear scaling factor, e.g., where Smis a positive scaling factor mapping the thrust commands Vthrustfrom the altitude controller 1113 to output thrust commands um1, um2

[0099] With the UAV 700 constantly rotating in the Z-direction, various example embodiments neglect the Z-axis so as to reduce the control to two axes via the calculation of its TPP, which is the plane perpendicular to the normal of the UAV 700 as illustrated in FIG. 11C, rz= [0, 0, 1]T, n = [nx, ny, nzY = Mrz(Equation 10) where rzdenotes the reference vector, which is a normal vector of the world plane and used as a reference to compare with the normal vector n of the TPP, and M denotes a matrix comprising the current attitude (quaternion) of the UAV 700 to be compared against the reference vector rz. By way of an example only and without limitation, M may be expressed as follows:(Equation 11)

[0100] The TPP pitch and roll angles may be extracted via the following operations:Ipitch utun2( ny, nz'), Goii atan2(nx,nz)'(Equation 12)

[0102] To simplify notation, various example embodiments introduce a function PID(sp, state) which refers to a PID controller with setpoint sp. and measuring state for feedback control. In this regard, in the R-mode attitude controller 1112, the TPP controller 1120 may comprise two independent PID controllers for the two axes (i.e., roll and pitch axes), respectively, which may be expressed as follows:(Equation 13) Therefore, according to Equation 13, a roll PID controller may be configured to receive the current roll state Γrolland the desired roll state setpoint Γroll , spand generate control outputsUcpg, xto reach the roll setpoint value Γroll, sp. Similarly, a pitch PID controller may be configured to receive the current pitch state Γpitchand the desired pitch state setpoint Γpitch, spand generate control outputs Ucpg, yto reach the pitch setpoint value Γpitch, sp.

[0102] Accordingly, the TPP controller 1120 may be configured to receive the TPP error and output the locomotion commands Ucpg, x, Ucpg,yto the CPG controller 1122. FIG. 11D depicts a schematic drawing of an example CPG controller 1122 according to various example embodiments of the present invention. As shown in FIG. 1 ID, the inputs to the CPG controller 1122 are Ucpg yand Ucpg xand an example differential equation relating them may be defined as follows:(Equation 14) where Φiand rirepresent the phase and amplitude of the i-th oscillator respectively, τrdenotes a positive gain which determines the speed which riconverges to the target amplitude Ri , and λiindicates the strength of the influence of the pacemaker oscillator on the i-th oscillator. In this regard, the pacemaker oscillator phase may be directly linked with the UAV heading, as shown in FIG. 1 ID. iq is the control output command for the i-th flap oscillator for controlling i-th flap 728a, 728b, where i = 1 or 2 which corresponds to the control output commandfor the first flap 728a and the control output command uf2for the second flap 728b, respectively.

[0103] Accordingly, the first control output commands comprise a first flap actuator output command (uf1) for controlling the first flap actuator 726a, a second flap actuator output command (uf2) for controlling the second flap actuator 726b, a first thrust output command (um1) for controlling the first thrust unit 724a and a second thrust output command (um2) for controlling the second thrust unit 724b. In this regard, the R-mode attitude controller 1112 is configured to generate the first and second flap actuator output commands (uf1uf2) of first control output commands for controlling the flight of the UAV 700 for the R-mode based on a pitch angle and a roll angle of a desired tip path plane of the UAV 700. In addition, the R-mode attitude controller 1112 is configured to generate the first and second thrust output commands (uml, um2) based on the thrust commands Vthrustreceived from the altitude controller 1113. By having the above TPP control, various example embodiments adopt a quadrotor-like attitude control for the pilot, where the radio control stick position directly controls the desired TPP pitch and roll angles, causing the UAV 700 to tilt and move in the commanded direction. Accordingly, various example embodiments advantageously reduce the control of the UAV 700 in the R-mode to a spinning disc and the number of rotational dimensions to 2. The spinning disc may also be expressed in the world frame J. As described above, the TPP pitch and roll angles may be referenced about the magnetic North direction.Quaternion-Based Attitude Controller (T / C Mode)

[0104] FIG. 1 IE depicts a schematic block diagram of the T / C-mode attitude controller 1114 according to various example embodiments of the present invention. Various example embodiments use the same attitude controller for both T-mode and C-mode, which is configured to have a cascaded control architecture comprising an inner angular rate controller (an inner loop for body rate angular control) and an outer quaternion attitude controller (an outer quaternion-based control loop which outputs the desired body rates). The desired quaternion may be inferred from the control signal from an operator input.

[0105] As shown in FIG. 11E, the outer loop of the T / C-mode attitude controller 1114 comprises a quaternion-based attitude controller 1130 configured to convert attitude quaternion setpoint commands (qsp) into body frame angular rate setpoint commands (ωsp). For example, the quaternion-based attitude controller 1130 may be derived from the nonlinear attitude control disclosed in D. Brescianini et al., “Nonlinear quadrocopter attitude control: Technical report, ETH Zurich 2013”. According to various example embodiments, qerrdenotes the error quaternion describing the difference between the desired quaternion qspand the currentquaternion of the UAV (q), i.e., qerr= q- 1 . qsp. In this regard, it will be understood by a person skilled in the art that, in quaternion terms, the difference between two quaternions is defined as the inverse of one quaternion multiplied with the other quaternion. In various example embodiments, an example control law may be expressed as follows:(Equation 15) where toω spdenotes to the desired angular rate setpoints and kp,qdenotes the proportional gain vector [kp q x, kp q y, kp q z], The sign() operator returns 1 for positive numbers and -1 for negative numbers. qerr,orefers to the real part of the quaternion (first element) and qerr,:13 refers to the imaginary part of the quaternion (last 3 elements), assuming that the quaternion is in wxyz format. Accordingly, the right hand side of Equation (15) dictates the values of the angular rate setpoints, based on the current state of the UAV attitude.

[0106] As the magnitude angular rate in the vertical Z-axis is large when transitioning from the R-mode, various example embodiments implement a reduced attitude control prioritizing pitch and roll stabilization over yaw. The amount of yaw priority may be controlled by the variable Pyaw, which may be slowly blended as the transition sequence progresses.

[0107] The angular rate controller 1132 may be configured to control the body frame angular velocity based on a series of PID controllers for each axis. The output of the angular rate controller 1132 may be virtual torque commands Vtin the body frame. In various example embodiments, the angular rate controller 1132 may be configured to produce the virtual torque commands VT (e.g., in the x-, y- and z-axes) based on the desired angular rate ωsp(desired angular rate setpoints) and the current angular rate to of the UAV 700 as follows:(Equation 16) In particular, the PID operation in Equation (16) returns a control output that attempts to set the the current angular rate ω to the desired angular rate ωsp, such as according to the following equation:(Equation 17)

[0108] In various example embodiments, unlike a typical tailsitter with counter-rotating propellers, the UAV 700 in the T-mode according to various example embodiments has bothpropellers (thrust units 724a, 724b) turning in the same direction, thereby generating a torque offset while hovering.

[0109] In various example embodiments, as shown in FIG. 1 IE, the T / C-mode attitude controller 1114 further comprises a control allocator 1134 configured to combine the virtual torque commands VT output from the angular rate controller 1132 and the virtual thrust commands Vthrustoutput from the altitude controller 1113 to form combined or concatenated virtual thrust / torque commands VTT (=[VTHRUST, Vt]) and map (or translate) the combined virtual thrust / torque commands VTT to the control output commands UTC(i.e., for the T-mode and the C-mode, e.g., comprisingUf2, um1, um2) for controlling the flight of the UAV 700. In this regard, various example embodiments employ the control allocator 1134 to derive the control output commands UTCwith appropriate weightage. In various example embodiments, a linear map of the virtual torque command VTT to the control output commands UTCmay be defined by a control allocation matrix Eg. Accordingly, the virtual torque commands VT output from the angular rate controller 1132 may be fed to the control allocator 1134, combined with the virtual thrust commands VTHRUSTto form the concatenated virtual thrust / torque commands VTT, to determine the control output commands UTCfor the thrust units 724a, 724b and the flap actuators 726a, 726b based on a control allocation matrix TB, which may be expressed as:UTc=ΓRVTT(Equation 18)

[0110] In various example embodiments, the control allocation matrix ΓBmay be determined by simulating a grid of feasible / suitable actuator output commands Usimand obtaining the simulated forces and torque (VTT sim) acting on the UAV 700. A least-squares linear fit may then be used to determine the control effectiveness matrix ΓG,which may be used to determine the control allocation matrix Eg via its pseudo-inverse, which may be expressed as follows:(Equation 19)(Equation 20) where ΓG+refers to the pseudo-inverse of ΓG.

[0111] The control allocation matrix ΓBmay be determined at multiple points of transition, and the matrices may then be interpolated element-wise according to the progress of the transition sequence. This interpolation only occurs during the R-mode to T-mode transition orthe T-mode to R-mode transition, as these particular transitions involve a mechanical change which changes the mapping of the virtual thrust / torque commands VTTto the control output commands UTC. In other words, the control allocation matrix interpolation is not required for transitions between the T-mode and the C-mode as they do not involve a mechanical change (i.e., the wings do not change angles) unlike transitions between the R-mode and the T-mode. Accordingly, the control allocation matrix Γ B may be determined via simulating the force and torque responses of the UAV 700 with respect to a large number of different actuator commands at various points of the transition.

[0112] Accordingly, the T / C-mode attitude controller 1114 is configured to generate second control output commands UTCfor controlling the flight of the UAV 700 for the T-mode and for the C-mode based on a desired quaternion attitude of the UAV 700. In this regard, as described above, the second control output commands UTCcomprise a first flap actuator output command uf1for controlling the first flap actuator 726a, a second flap actuator output command Uf2for controlling the second flap actuator 726b, a first thrust output command um1for controlling the first thrust unit 724a and a second thrust output command um2for controlling the second thrust unit 724b. In various example embodiments, as described above, the T / C-mode attitude controller 1114 is configured to: generate virtual torque commands Vt based on the desired quaternion attitude of the UAV 700; and generate the second control output commands UTCbased on the virtual torque commands VT and a control allocation matrix Γ B. In various example embodiments, the virtual torque commands VT may be combined with the virtual thrust commands Vthrustto form combined virtual thrust / torque commands VTT and the control allocation matrix Γ B may be configured to map the virtual thrust / torque commands VTT to the second control output commands UTC.Blending Controllers

[0113] A simplest way to switch between different controllers for different flight modes during a transition between flight modes may be to simply switch the routing for controlling the UAV from the control output commands generated by one controller to those of another controller, akin to an ‘on-off or a hard switch. In contrast, various example embodiments seek to provide a transition control which optimizes conditions (e.g., the UAV state) for an incoming controller for an incoming flight mode to take over from an outgoing controller for an outgoing flight mode, while maintaining the stability and control needed to keep the UAV 700 in the air. In this regard, as the R-mode and the T / C-mode utilize fundamentally different controllers,various example embodiments provide a simple and effective blending strategy or technique between these different controllers for different flight modes, gradually handing off control of the thrust units 724a, 724b and the flap actuators 726a, 726b from one controller to another controller during a transition, instead of a hard switch between them. As will be described later below, various example embodiments further optimize various blending or transition parameters (e.g., blending points and the speed of the blending) to enhance flight stability and effectiveness.

[0114] As described hereinbefore, the flight controller 1160 comprises a mode transition controller 1116 comprising the control blending block 1117 communicatively coupled to the R-mode attitude controller 1112 and the T / C-mode attitude controller 1114 for mixing or blending the control output commands UR, UTCtherefrom during a transition from one flight mode to another flight mode. In this regard, the control blending block 1117 may be configured to mix or blend the control output commands UR, UTCfrom the R-mode attitude controller 1112 and the T / C-mode attitude controller 1114 based on a transition coordinator variable or parameter, such as a wing angle variable or parameter ( Bangle). In various example embodiments, the mode transition controller 1116 may further comprise the C-transition control block 1118 configured to, for a transition between the T-mode and the C-mode, modify the target base pitch angle associated with or provided by the desired quaternion attitude prior to the desired quaternion attitude (as modified) being used by the T / C-mode attitude controller 1114 to generate the control output commands. In particular, the C-transition control block 1118 may be configured to control an amount of target base pitch angle modulation (by directly modifying the desired quaternion attitude) based on a state of the transition (Bcruise) between the T-mode and the C-mode For example, Bangle,BcruiseG [0, 1] , where Bangle= 0 corresponds to the wing actuator 730 (and thus the wings 720a, 720b) being in the extended state in the R-mode and Bangle= 1 corresponds to the wing actuator 730 (and thus the wings 720a, 720b) being in the retracted state in the T / C-mode. Bcruise= 0 corresponds to the UAV 700 being fully set to operate in the R-mode or the T-mode (i.e., not in the C-mode), while Bcruise=1 corresponds to the UAV 700 being fully set to operate in the C-mode. As explained hereinbefore, in various example embodiments, the T-mode and the C-mode utilize the same T / C-mode attitude controller 1114 (inner angular rate control loop cascaded with an outer quaternion attitude control loop), and the difference being the base pitch target (target base pitch angle) for the UAV 700. In particular, in the T-mode, the base pitch target may be set such that the wings 720a, 720b are oriented perpendicular to the ground plane during flight, with thethrust units 724a, 724b pointing up (vertically). In the C-mode, the base pitch target is set such that the UAV 700 is pointing in a forward direction, with the wings 720a, 720b oriented parallel to the ground, like a fixed-wing aircraft.

[0115] In various example embodiments, a plurality or series of control weights, br, bt , and be, corresponding to the R-mode, the T-mode and the C-mode, respectively, may be provided to control (based on br and bt) how much mixing of the control output commands UR, UTCfrom the R-mode attitude controller 1112 and the T / C-mode attitude controller 1114 to produce the main or final control output commands Umainfor controlling the flight of the UAV 700 and to control (based on be) the amount of target base pitch angle modulation for modifying the target base pitch angle associated with or provided by the desired quaternion attitude prior to being used by the T / C-mode attitude controller 1114 to generate the control output commands UTC. Control weights br, bt , and be may also be referred to as R-control weight (or R-transition control weight), T-control weight (or T-transition control weight) and C-control weight (or C- transition control weight), respectively. Accordingly, control blending weights br and bt control the amount of the first control output commands URand the amount of the second control output commands UTCfor contributing in the resultant control output commands Umain(e.g., for forming or making up the final control output commands) for controlling the flight of the UAV 700. In various example embodiments, each of the plurality of control weights may be expressed as a control function of a transition coordinator variable, Bangleor Bcruise(e.g., for control blending weights br and bt, a blending function of transition coordinator variable Bangle), of which may be normalized. In various example embodiments, the characteristics of the control for a transition may be configured or adjusted by parameterizing the control function as a sigmoid equation.

[0116] Accordingly, the control blending block 1117 is communicatively coupled to the R- mode attitude controller 1112 and the T / C-mode attitude controller 1114 and configured to, for a transition between the R-mode and the T-mode, control an amount of the first control output commands URand an amount of the second control output commands UTCfor contributing in resultant control output commands Umain(e.g., final control output commands) for controlling the flight of the UAV 700 based a state of the transition between the R-mode and the T-mode. In this regard, for the transition between the R-mode and the T-mode, the amount of the first control output commands URand the amount of the second control output commands UTCfor contributing in the resultant control output commands Umainare determined based on a wing angle parameter Bangle(and thereby the values of control blending weights br and bt). In thisregard, the wing angle parameter Banglecorresponds to a collective angle of the wing members 720a, 720b and represents the state of the transition between the R-mode and the T-mode. Furthermore, for a transition from the R-mode to the T-mode, the amount of the first control output commands URand the amount of the second control output commands UTCfor contributing in the resultant control output commands Umainare determined based on a first sigmoid function (based on which control blending weights and bt for the R-mode to the T- mode transition are determined) configured as a function of the wing angle parameter Bangle. Similarly, for a transition from the T-mode to the R-mode, the amount of the first control output commands URand the amount of the second rotorcraft control output commands UTCfor contributing in the resultant control output commands Umainare determined based on a second sigmoid function (based on which control blending weights br and bt for the T-mode to the R- mode transition are determined) configured as a function of the wing angle parameter Bangle.

[0117] Accordingly, in various example embodiments, the C-transition control block 1118 is configured to, for a transition between the T-mode and the C-mode, control an amount of target base pitch angle modulation based a state of the transition between the T-mode and the C-mode (corresponding to the Bcruisevariable or parameter, and thus, based on the value of C- control weight be). In various example embodiments, for a transition from the T-mode to the C-mode, the amount of target base pitch angle modulation is determined based on a third sigmoid function (based on which C-control weight be for the T-mode to the C-mode transition is determined) configured as a function of the state of the transition between the T-mode and the C-mode (corresponding to the Bcruisevariable or parameter). In various example embodiments, for a transition from the C-mode to the T-mode, the amount of target base pitch angle modulation is determined based on a fourth sigmoid function (based on which control target pitch weight be for the C-mode to the T-mode transition is determined) configured as a function of the state of the transition between the T-mode and the C-mode. In various example embodiments, C-control weight be may directly control the target base pitch angle by a simple scaling method according to an example equation: new_desired_target_base_pitch_angle = bc* desired_target_base_pitch_angle Accordingly, the C-transition control block 1118 may be configured to directly modify the desired quaternion attitude according to the C-control weight be to perform the target base pitch angle modulation.

[0118] In various example embodiments, the shape of a control function (e.g., the above- mentioned first, second, third or fourth sigmoid function) associated with a transition may be defined or characterised by function parameters as follows:• transition midpoint parameter TFT.C and transition width parameter TFT, W - provided to control the center and the width of the control (e.g., blending) function (e.g., the above- mentioned first sigmoid function) for blending control output commands UR, UTCfrom the R-mode attitude controller 1112 and the T / C-mode attitude controller 1114 for a transition from the R-mode to the T-mode;• transition midpoint parameter TBT.C and transition width parameter TBT, W - provided to control the center and the width of the control (e.g., blending) function (e.g., the above- mentioned second sigmoid function) for blending control output commands UR, UTCfrom the R-mode attitude controller 1112 and the T / C-mode attitude controller 1114 for a transition from the T-mode to the R-mode;• transition midpoint parameter TFc.c and transition width parameter TFc, w - provided to control the center and the width of the control function (e.g., the above-mentioned third sigmoid function) controlling the amount of target base pitch angle modulation for a transition from the T-mode to the C-mode; and• transition midpoint parameter TBc.c and transition width parameter TBc, w - provided to control the center and the width of the control function (e.g., the above-mentioned fourth sigmoid function) controlling the amount of target base pitch angle modulation for a transition from the C-mode to the T-mode.For example, the above transition width parameter directly adjusts the horizontal scaling of the inflection point of the sigmoid curve defined by Equations (21) and (22) to be described below. When the scale of the inflection point is larger, the slope / steepness of the sigmoid curve becomes gentler. Subsequently, when the scale of the inflection point is smaller, the sigmoid curve becomes steeper.

[0119] In various example embodiments, there is provided an additional function parameter TRNc.t which defines the time taken for the T-mode to C-mode transition. For example, the blending time for Bangle may be fixed at 6.5 seconds which is the time taken for the wing actuator 730 to reach from one endpoint to the other. Regarding the above-mentioned function parameters, those starting with TF correspond to a forward transition direction (e.g., R-mode to T-mode to C-mode, or simply referred to as R-T-C) and those starting with TB correspond to abackward transition direction (e.g., C-mode to T-mode to R-mode, or simply referred to as C- T-R).

[0120] For illustration purpose, FIGs. 12A and 12B depict a plot of control weights (sigmoid control weights, namely, br, bt and bc) for the forward transition direction and the backward transition direction, respectively, and illustrates the effect of the above-mentioned function parameters on the shape of the control weights. By way of examples only and without limitations, the control weights br, bt, and be, for the forward transition direction and the backward transition direction may be expressed as follows:Forward Transition Direction (R-T-C):(Equation 21) Backward Transition Direction (C-T-R):(Equation 22)

[0121] For example, bt = 1 corresponds to the completion of the transition from the R-mode to the T-mode, and be = 1 corresponds to the completion of the transition from the T-mode to the C-mode. Accordingly, when bt = 1 and be = 1, this indicates that the entire forward transition from the R-mode to the T-mode (R-T-C) has been completed, whereby the adjustment of the base pitch angle of the UAV 700 to the C-mode has been completed.

[0122] Accordingly, in various example embodiments, the above-mentioned first sigmoid function is parameterized based on one or more first function parameters comprising one or more of a transition midpoint parameter and a transition width parameter for the first sigmoid function. In various example embodiments, similarly, the above-mentioned second sigmoid function is parameterized based on one or more second function parameters comprising one or more of a transition midpoint parameter and a transition width parameter for the second sigmoid function. In various example embodiments, the above-mentioned third sigmoid function isparameterized based on one or more third function parameters comprising one or more of a transition midpoint parameter and a transition width parameter for the third sigmoid function. In various embodiments, similarly, the fourth sigmoid function is parameterized based on one or more fourth function parameters comprising one or more of a transition midpoint parameter and a transition width parameter for the fourth sigmoid function.

[0123] For illustration purpose, FIG. 13 shows plots of the blending of control output commands from the R-mode attitude controller 1112 and the T / C-mode attitude controller 1114 to produce the final control output commands for controlling the flight of the UAV 700, based on the angle of the wings 720a, 720b (Bangle) and the fixed-wing transition state (Bcruise). From the plots, it can be observed how the control weights (br,bt and bc) change as the angle of the wings 720a, 720b (Bangle) changes, followed by the state of the C-mode transition cruise)-

[0124] Various factors contributing to altitude changes during R-to-T and T-to-R transitions will now be discussed according to various embodiments of the present invention.

[0125] During the R-to-T transition, it is desirable for the UAV 700 to stop rotating, and during the T-to-R transition, it is desirable that the UAV 700 spins up as fast as possible to generate the required lift as the angle of the wings 720a, 720b change. This can be achieved via the flaps 728a, 728b on the UAV 700. However, due to the nature of the UAV 700, these flap movements may cause unintended changes in altitude.

[0126] For the R-to-T transition, the flaps 728a, 728b are controlled to point downward, effectively increasing the amount of lift generated and resulting in an increase in altitude. On the other hand, during the T-to-R transition, the flaps 728a, 728b point upward, reducing the lift force and resulting in a drop in altitude. These changes in altitude can be compensated via increasing or decreasing the throttle respectively, however, varying the throttle may also in turn vary the effectiveness of the flaps 728a, 728b. Therefore, in various example embodiments, a balance is set between the throttle used and the allowed altitude deviation. In particular, according to various example embodiments, the function parameters are optimized to achieve various desired performance goals or requirements, which will now be described below.Optimization of function Parameters

[0127] As described hereinbefore according to various example embodiments, a number of function parameters may be provided to configure or characterize the shape of a control function (e.g., the above-mentioned first, second, third or fourth sigmoid function) associated with atransition, such as but not limited to,TFT C, TFT W, TFC C, TFC W, TBT C, TBT W, TBC C, TBC W, Tc t. In various example embodiments, to improve the transition control sequence between flight modes (e.g., compared to manually adjusting the function parameters via trial and error), an optimization process is performed to determine the values of these function parameters. For example, the optimization may be performed offline in simulation before being implemented on the UAV 700. For example, the function parameters being optimized may be represented by an optimization vector E, where s. = [TFT C, TFT W, TFc c, TFc w, TBT C, TBT W, TBc c, TBc w, Tc t]T. In this regard, FIG. 14 shows a table (Table 1) of example function parameters which may be optimized according to various example embodiments of the present invention.

[0128] In the optimization, according to various example embodiments, a complete simulation routine may be defined as an ‘episode’, with the episode returning a single value defining the ‘fitness’ of the input solution. For such episodic optimization problems, for example, Parameter-Exploring Policy Gradients (PEPG) may be utilized for the optimization algorithm, which is an Evolutionary Algorithm (EA).

[0129] For the optimization problem, according to various example embodiments, an objective function is defined to be maximized. In the objective function, a constant may be included per timestep as an incentive for the optimization to favor longer episode lengths. For example, angular velocity, cartesian velocity, control outputs, and the control output derivatives may be added as a cost to be minimized. An additional ‘completion bonus’ may be added should the episode run from start to the maximum specified end time of 8 second (he™ = 400), where iterm denotes the last timestep just before the episode terminates. There may be two sets of simulations per episode, the first episode is whereby the UAV 700 is brought through the forward transition sequence from the R-mode to the C-mode, passing through the T-mode in the process, and the second episode is whereby the UAV 700 is brought through the back transition sequence from the C-mode to the R-mode, also passing through the T-mode in the process. A status parameter Tdir denotes the current transition direction of the simulation, Tdir= 1 denotes that the transition is in the forward transition direction (R-T-C direction) and Tdir= — 1 denotes that the transition is in the backward transition direction (C-T-R direction). Two additional terms may be provided in JT specific to the forward transition direction (Tdir= 1), the first term being an added cost to the end throttle and the second term being a maximization of the distance in the Y-direction to incentivize the UAV 700 to transit to forward flight. For example, an example objective function may be expressed as follows,(Equation 23)

[0130] In Equation (23), JPrefers to the sum of the penalty functions, such as defined inTable 2 shown in FIG. 15. The term Jcdenotes the ‘completion bonus’ if the last timestep (iterm) is 400:(Equation 24)

[0131] The term JT denotes the sub-objective function that changes depending on the transition direction:(Equation 25)

[0132] The term ji is applied every timestep and summed up across the entire episode:(Equation 26)

[0133] FIG. 16 depicts a schematic flow diagram of the optimization process associated with each episode to obtain the objective function. Each episode may include two subsimulations for both R-to-C direction and C-to-R direction. As shown in FIG. 16, the episode begins by checking if the function parameters suggested by the optimizer is feasible, if not feasible, the episode ends and returns the penalty function defined in Table 2 shown in FIG 15. If feasible, the simulation from the R-to-C direction begins. The simulation is advanced by 1 timestep until a stop criteria is met. The simulation from the C-to-R direction then begins, and similarly, the simulation is advanced by 1 timestep until the stop criteria is met. Based on the data logged during the simulation, the simulation cost function is calculated. The value of the simulation cost function is then returned and the episode ends.

[0134] Accordingly, in various example embodiments, the above-mentioned one or more first function parameters, the above-mentioned one or more second function parameters, the above-mentioned one or more third function parameters and the above-mentioned one or morefourth function parameters are optimized (e.g., collectively optimized) based on the objective function.

[0135] In various example embodiments, the overall penalty function, Jp, is returned if the optimization vector is not feasible. Feasibility constraints are set on the function parameters to keep the operating range to within 0 and 1, corresponding to the normalized transition coordinator variables Bangle and Bcruise. This is to prevent scenarios such as the blending taking longer than the transition coordinator allows, which would result in the control weights br, bt, and be not reaching their endpoints. The transition time for the C-mode is also constrained to a positive value. Quadratic penalty functions are imposed and are simply added to the objective function if any of the constraints are broken. This would enable the algorithm to automatically learn the feasible set as the optimization progresses. As mentioned above, Table 2 shown in FIG. 15 shows the constraints of the function parameters and their respective penalty functions.

[0136] Regarding the initial conditions for the R-T-C direction, that is, in the case where the simulation is in the forward transition direction (Tdir = 1), the UAV is placed at the origin, with the yaw (Z-axis) rotation uniformly randomized between 0 to 2πt (360 degrees). Initial angular velocity in all axes are also uniformly randomized between -0.5 and 0.5 rad / s. With an additional 20 rad / s added to the Z-axis (body frame).

[0137] Regarding the initial conditions for the C-T-R direction, that is, in the case where the simulation is in the backward transition direction (Tdir = -1), the UAV is placed at -10 m from the origin in the Y-direction. With the yaw (Z-axis) at zero. Initial angular velocity in all axes are also uniformly randomized between -0.5 and 0.5 rad / s. The initial pitch angle is offset by π / 8 rad to simulate the UAV being initialized in the C-mode.

[0138] Regarding episode termination conditions, each episode is subject to termination constraints where a violation would result in the termination of the episode, directly affecting the objective function value. This way, the optimization would be incentivized to avoid solutions that would place the UAV in such states, such as drifting too far and wobbling too much during the transition sequence. An example list of termination constraints is: (Restrict x and z axes to a bounding range of 1 m each side),(Angular velocity (x and y axis) magnitude bounding range), and• t = 8s (Maximum episode time. At a timestep of 0.02s, this corresponds to 400 timesteps).

[0139] The optimization was run for a total of 420 generations, with a population size of 100, initial standard deviation of 0.25, and a learning rate of 0.1. The solution for the functionparameter values returned by the optimization process is presented in Table 3 in FIG. 17. FIG. 18 depict plots of the values of the control weights (br,bt , bc) in relation to the transition coordinator variables (Bangle, Bcruise) for a switched controller (top of FIG. 18) and the optimized blended controller (bottom of FIG. 18). Over a total of 5 consecutive runs, the maximum objective function value reached was 488.9304. The progression of the objective function value with the iterations is presented in FIG. 19. In particular, FIG. 19 depicts a plot of the fitness value throughout the optimization according to various example embodiments of the present invention.Simulation and Analysis

[0140] A simulation analysis which uses the dynamic model described hereinbefore will now be presented according to various example embodiments of the present invention. In this regard, the simulation analysis is focused more on the R-mode and the T-mode transitions. In the simulation analysis, a comparison was made between a simple switched controller where the controllers are instantly switched at a fixed point in the transition and the optimized sigmoid blending control strategy according to various example embodiments of the present invention.

[0141] FIGs. 20A and 20B show the transition simulation data obtained. FIG. 20A shows a plot of the XY position histories of 10 consecutive simulations from R-T-C modes, akin to a top view. The plots present the transition simulation data for unoptimized switched control transitions and optimized blended control transitions. In particular, FIG. 20A presents the X- and Y-axis deviation as the UAV undergoes the stated transition sequence. The optimized transition displays less position deviation compared to the non-optimized transitions, with a mean absolute value of 0.0993 m for the optimized transition and 0.2905 m for the unoptimized transition, presenting an improvement of 65.8% and indicating that the optimized transition exhibits more stable and smoother behavior. For the C-T-R direction, the UAV was set to initially approach from the negative Y-axis direction. Visually, in both directions, R-T-C and C-T-R, the optimized controller exhibits less position deviation and an overall smoother and more predictable transition.

[0142] With respect to the throttle commands (Um1, Um2) for the thrust units, FIG. 20B shows the simulated throttle comparison of the R-mode to the T-mode transition (R-to-T transition) and of the C-mode to R-mode transition (C-T-R transition) for the unoptimized switched control and optimized blended transitions. The throttle represented in the graphs in FIG. 20B (as well as FIG. 20C) is the sum of both motors (Um1and Um2). It can be observedthat the non-optimized transition has a significant throttle spike in the middle of the transition (at a maximum of around 1.25), which would translate to greater power draw and more potential for undesirable movements. On the other hand, the optimized transition is more desirable and results in a significantly smoother transition between the flight modes.Experimental Evaluation

[0143] For illustration and experimental purposes, FIGs. 21A and 21B depict photos of a UAV prototype used in experiments conducted, corresponding to the UAV 700 as described hereinbefore according to various embodiments, in the R-mode and in the T / C-mode, respectively. The UAV comprises two flat wooden wings attached to a 3D-printed central hub which houses the transition mechanism and the control electronics. For example, flight data may be logged onboard using a flash memory card, such as an SD card. The UAV prototype further comprises an onboard GPS receiver to log the position data during the outdoor field tests. In experiments conducted, the performance of the optimized transition parameters is evaluated to demonstrate the full multi-modal capabilities of the UAV prototype according to various example embodiments of the present invention. Experiments were conducted in a controlled lab environment with motion capture assistance, as well as in outdoor experiments in real world conditions.

[0144] For the UAV prototype, onboard state estimation is provided using an onboard 9- axis IMU (Invensense ICM-20948). Attitude quaternion is provided by the sensor’s Digital Motion Processor (DMP) and the angular rate is determined directly from the gyroscopic sensor. Attitude and angular rate data is polled at 200 Hz. To verify the quality of the state estimation provided by the DMP, several test flights were conducted and the attitude state estimation data were compared with a motion capture ground truth. The main flight computer is an ESP32 Pico board microcontroller running FreeRTOS comprising two Xtensa LX6 CPUs. Real-time telemetry is transmitted using its built in 2.4 GHz WiFi to a ground control device utilizing the MAVLink protocol. FIG. 22 depicts a schematic block diagram of various electronics components onboard the UAV prototype and offboard, as well as showing example communication paths (coupling / connections) between various electronics components. The propulsion system comprises two thrust units, each implemented by a Tiger Motor F40 with 1600 RPM / V (Kv) paired with a DJI Triblade 5045 propeller (5 inch diameter with 4.5 inch blade pitch). Each thrust unit provides a maximum of 6.4 N (650 g) of thrust when operated with a 4S Lithium Polymer (LiPO) battery (14.8 V nominal voltage). A L12-R Micro LinearServo with a stroke length of 30 mm and gear ratio 210: 1 was used for controlling the collective angle of the two wings. This linear servo was selected for its minimal backlash and high actuation torque. A laser-cut wooden flat plate wing was used for ease of fabrication. The wing flaps are driven directly by flap servos (Hitec 5065MG+) in a 3D-printed housing beside the wing flaps. By way of an example only and without limitations, various physical properties of the prototype UAV are presented in Table 4 shown in FIG. 23 according to various example embodiments of the present invention. For example, the ESP32 Pico receives user commands via a serial connection from the Spektrum Serial receiver, which receives the user commands transmitted by a radio controller held by the user. The ESP32 Pico also interfaces via SPI to an SD Card Logging Module for datalogging purposes. The ESP32 Pico generates PWM signals for commanding the transition actuator (the linear actuator), the flap actuators / servos and the electronic speed controller (and therefore the thrust units).

[0145] Firstly, flight tests in a controlled environment were conducted to demonstrate the repeatability of the UAV prototype to transit between the R-mode and the T-mode (R-to-T and T-to-R transitions) over a period of 140 seconds. FIGs. 24A to 24C show plots of the UAV flight data obtained during the R-to-T and T-to-R transitions. In particular, FIG. 24A shows a plot of the angular velocities in the body frame during the repeated R-T-R transitions which were invoked manually by a user. In the plot, the lighter grey background shading indicates the R-mode and the darker background shading indicates the T-mode. From the plot shown in FIG. 24A, the most prominent change in angular velocity can be observed in the z-axis, being almost stationary during T-mode and reaching the operating range of around 25-30 rad / s in the R-mode, and gaining or losing angular velocity as required. A total of six consecutive transitions were conducted while the UAV prototype was flying.

[0146] The three different transition controls are also compared, namely, the basic switched control, the baseline blended control and the optimized blended control (optimized according to various example embodiments of the present invention). The metrics used include the altitude deviation, as well as the throttle output during the transitions. The transitions for each controller were conducted consecutively and repeated for four times. FIG. 24B show plots of the altitude data of the R-to-T transitions and T-to-R transitions of the three transition controls, the plot for each transition control shows the mean of four flights with the bounding area representing 1 SD (standard deviation). The mean absolute error of the altitude is also shown in FIG. 24B. It can be observed that the optimized blended control deviates the least from the desired altitude. FIG. 24C shows plots of the summed throttle data of the R-to-T transitions and T-to-R transitions ofthe three transition controllers, the plot for each transition control showing the mean of four flights with the bounding area representing 1 SD. It can be observed that the optimized blended control exhibits less sudden spikes and overall utilizes less throttle to complete the transitions. Accordingly, from the plots, it can be observed that the optimized blended control performs the best in terms of both altitude deviation and throttle usage. In contrast, the switched transition control exhibits the worst performance. Furthermore, it can be observed that implementing a basic but unoptimized blended transition control (baseline blended control) improves the altitude deviation by 30% during R-to-T transition, and 20.2 % during T-to-R transition. The optimized blending control according to various example embodiments of the present inventio further reduces it by 14.7 % for R-to-T transition, and by 12.9 % for the T-to-R transition. The results are summarized in Table 5 in FIG. 25, which presents the Altitude Mean Absolute Error (MAE) and Throttle Sum (normalized) during the R-to-T mode transitions as well as T-to-R mode transitions.

[0147] In addition, numerous flight tests outdoors in real-world conditions were also conducted mainly to evaluate the multi-modal flight capabilities of the UAV prototype. All outdoor flight tests were performed in an open field. FIG. 26A depicts plots of mean XY position history for each of the three different transition controls being compared, where the desired direction to transition to the C-mode is in the positive Y-axis direction. FIG. 26B depicts plots of the throttle usage comparisons of the three different transition controls.

[0148] The flights conducted demonstrate utilizing all three flight modes in a single flight (three-mode flight capability), starting out in the T-mode via hand-launch, transiting into the R-mode, back to the T-mode, cruising in the C-mode, and transiting back to the T-mode. FIG. 27A presents this flight data in a map overlay of the test site. In particular, FIG. 27A show a plot of the UAV position history in a map overlay of the test site for a flight. FIG. 27B show a plot of the UAV position (in meters) against time and a plot of the UAV altitude against time. This particular flight was conducted in moderate wind conditions of around 5-20 km / h. A maximum altitude of around 20m above ground level was reached while the UAV was in the C-mode. The size of the testing space was approximately 150 m by 150 m.

[0149] In a separate flight, a top-down aerial view of the transitions were also captured from a separate drone. FIG. 28A shows a composite top-down aerial view of the transition process captured from another drone, which depicts the motion of the UAV prototype as it transitions between the three different flight modes (T-mode, R-mode and C-mode). Snapshots of the UAV prototype during the transition are also shown in FIG. 28A This flight wasconducted in moderate wind conditions of 5-20 km / h. The UAV prototype entered a transition from the T-mode to the R-mode at an altitude of about 3 m, followed by a R to T mode transition at the same altitude, before transiting to the C-mode and climbing to a maximum altitude of 20 m. The UAV prototype circled around the field 3 times before descending and transiting back to the T-mode, followed by a transition to the R-mode and landing. FIG. 28B shows the position history of the UAV prototype during a hairpin-like maneuver.

[0150] Voltage and current data were logged throughout the flight and multiplied to obtain the instantaneous power draw. The onboard sensors were pre-calibrated with a known source beforehand. They were then categorized into the respective mode of operation. FIG. 28C shows a boxplot of the power consumption during each of the three flight modes measured from onboard current and voltage sensors. From the boxplot shown in FIG. 28C, it can clearly be seen that the UAV prototype in the T-mode draws the most power, a mean of 140 W. The mean power consumption of the R-mode and the C-mode are relatively close at 96.9 W vs 114.4 W. This translates to 30% less power draw on average while in the R-mode. This is most likely due to the R-mode and C-mode utilizing the wings for lift instead of the propulsiox n system directly, which has a higher disc loading due to the smaller area than the wings.

[0151] It can also be observed that the variance of the power consumption during the C- mode is higher compared to that of the R-mode. This may be due to the nature of the fixed- wing flight, it being constantly moving at a forward velocity causes it to be more sensitive to the varying wind conditions, such as flying within a tailwind condition would require more thrust from the motors to maintain airspeed, compared to flying within a headwind condition which would require less thrust. It can also be noted that during windy conditions when flying in the R-mode, the UAV prototype tends to climb more easily, irrespective of wind direction, which implies an increase in lift whenever a wind gust is experienced.

[0152] One of the limitations of the R-mode is that the control flaps need to rotate the full round for the cyclical control, which adds delay from the control input to the UAV, compared to the more immediate actuation of the motors / propeller thrust in the T-mode. Accordingly, having multiple flight modes provides an option to a user if agility is desired, thus providing capability for agile maneuvers. While in the C-mode, the UAV is able to perform agile maneuvers during forward flight, such as a hairpin-like turn atypical of fixed-wing aircraft. This is due to the dual-motor (dual thrust units) configuration of the UAV according to various example embodiments, which provides differential thrust control, and the high thrust-weight ratio of the propulsion system, providing enough lift to keep the UAV in the air even at lowhorizontal velocities where the wings generate minimal lift. For example, this is demonstrated in the flight conducted as described above with reference to FIGs. 28A to 28C.

[0153] Accordingly, various example embodiments provide the development, construction and experimental evaluation of an agile multi-modal UAV capable of operating in three distinct flight modes (namely, the R-mode, the T-mode and the C-mode), and demonstrated the capability to seamlessly transit between the flight modes mid-air. A control-blending transition strategy or technique between all flight modes is also provided, optimized and tested on a physical prototype. The optimized blending controller according to various example embodiments was advantageously found to result in less altitude variation, especially in R-to- T and T-to-R transitions, while also utilizing less throttle compared to a basic switched transition control. In various example embodiments, the outdoor flight control may be implemented in semi-auto mode with attitude stabilization or may be implemented without reliance on motion capture for position control. By implementing flight control without reliance on motion capture, fully autonomous flights and transitions can be conducted without any input from a pilot / user.

[0154] Accordingly, various example embodiments advantageously provide a multi-modal, nature-inspired UAV that can operate in three different flight regimes (rotating rotorcraft flight mode, non-rotating rotorcraft flight mode and fixed-wing flight mode), for example, unlike typical hybrid UAVs which only consist of two flight modes. The platform uses a rotating rotorcraft flight mode to achieve efficient hover, yet gives the flexibility to the user to enter a more agile hover state with an additional non-rotating rotorcraft flight mode. Both the mechanical configuration and software / control architecture used to achieve this three-mode capability have been described in detail herein according to various example embodiments of the present invention. A sigmoid blending control has been implemented according to various example embodiments as a transition control strategy, involving transition coordinators that adjust the weight of each individual controller (e.g., the R-mode attitude controller 1112 and the T / C-mode attitude controller 1114) on the control output commands (e.g., comprising a first flap actuator output commandfor controlling a first flap actuator, a second flap actuator output command Uf2for controlling a second flap actuator, a first thrust output command um1for controlling a first thrust unit and a second thrust output command um2for controlling a second thrust unit) accordingly. To improve the transition sequence based on desired performance goals such as reduced altitude variation and throttle usage, an optimization routine is also performed according to various example embodiments to obtain the optimized functionparameters. As discussed above, the optimized function parameters have been experimentally verified on a physical prototype in lab conditions and shown to exhibit at least a 20% improvement for altitude variation and at least 5% less throttle usage across 4 consecutive transitions compared to a baseline. Three-mode capability outdoors in real-world conditions was also demonstrated in multiple flights. Power draw in the rotating rotorcraft flight mode is found to be 30% less than in non-rotating rotorcraft flight mode.

[0155] While embodiments of the invention have been particularly shown and described with reference to specific embodiments, it should be understood by those skilled in the art that various changes in form and detail may be made therein without departing from the scope of the invention as defined by the appended claims. The scope of the invention is thus indicated by the appended claims and all changes which come within the meaning and range of equivalency of the claims are therefore intended to be embraced.

Claims

CLAIMS1. A hybrid aerial vehicle with multiple flight modes, comprising: a central member; a first wing member arranged at a first side of the central member, the first wing member having a first thrust unit and a first flap actuator affixed thereto, the first flap actuator having a first flap member affixed thereto and is configured to actuate the first flap member; a second wing member arranged at a second side of the central member, the second wing member having a second thrust unit and a second flap actuator affixed thereto, the second flap actuator having a second flap member affixed thereto and is configured to actuate the second flap member, wherein the first and second sides of the central member are opposite sides of the central member; a wing actuator configured as a linear actuator and configured to rotate the first and second wing members about a lateral axis of the central member via a first mechanical linkage and a second mechanical linkage, respectively; and a flight controller communicatively coupled to the wing actuator, the first and second thrust units and the first and second flap actuators and configured to control, in response to a transition control signal received, the hybrid aerial vehicle to transition from one flight mode to another flight mode amongst the multiple flight modes during flight based on controlling the wing actuator, the first and second thrust units and the first and second flap actuators, the multiple flight modes comprising a rotating rotorcraft flight mode, a non-rotating rotorcraft flight mode and a fixed-wing flight mode.

2. The hybrid aerial vehicle according to claim 1, wherein the first and second mechanical linkages are coupled to the first and second wing members, respectively, and are each further coupled to the wing actuator, and the first and second mechanical linkages are each configured to transform a linear motion of the wing actuator to a rotational motion for rotating the corresponding wing member of the first and second wing members about the lateral axis of the central member.

3. The hybrid aerial vehicle according to claim 2, wherein the wing actuator is configured to rotate the first and second wing members about the lateral axis of the central member in opposite directions via the first and second mechanical linkages, respectively.

4. The hybrid aerial vehicle according to any one of claims 1 to 3, wherein the first thrust unit and the first flap member are arranged at a leading edge and a trailing edge of the first wing member, respectively, and the second thrust unit and the second flap member are arranged at a leading edge and a trailing edge of the second wing member, respectively.

5. The hybrid aerial vehicle according to any one of claims 1 to 4, further comprising: a frame member arranged along the lateral axis of the central member, wherein the central member is coupled to the frame member, and the first and second wing members are rotatably coupled to the frame member such that the first and second wing members are rotatable about the frame member when actuated by the wing actuator.

6. The hybrid aerial vehicle according to claim 5, wherein the frame member is a rod.

7. The hybrid aerial vehicle according to any one of claims 1 to 6, wherein the central member is configured to house the wing actuator and the flight controller.

8. The hybrid aerial vehicle according to any one of claims 1 to 7, wherein the flight controller comprises a wing controller configured to: control, in response to the transition control signal being for transitioning the hybrid aerial vehicle from the rotating rotorcraft flight mode to the non-rotating rotorcraft flight mode, the wing actuator to rotate the first and second wing members about the lateral axis of the central member from an extended state to a retracted state, and control, in response to the transition control signal being for transitioning the hybrid aerial vehicle from the non-rotating rotorcraft flight mode to the rotating rotorcraft flight mode, the wing actuator to rotate the first and second wing members about the lateral axis of the central member from the retracted state to the extended state.

9. The hybrid aerial vehicle according to claim 8, wherein the flight controller further comprises:a first flight mode controller configured to output first control output commands for controlling a flight of the hybrid aerial vehicle for the rotating rotorcraft flight mode; a second flight mode controller configured to output second control output commands for controlling a flight of the hybrid aerial vehicle for the non-rotating rotorcraft flight mode; and a mode transition controller communicatively coupled to the first and second flight mode controllers and configured to, for a transition between the rotating rotorcraft flight mode and the non-rotating rotorcraft flight mode, control an amount of the first control output commands and an amount of the second control output commands for contributing in resultant control output commands for controlling the flight of the hybrid aerial vehicle based a state of the transition between the rotating rotorcraft flight mode and the non-rotating rotorcraft flight mode.

10. The hybrid aerial vehicle according to claim 9, wherein, for the transition between the rotating rotorcraft flight mode and the non-rotating rotorcraft flight mode, the amount of the first control output commands and the amount of the second control output commands for contributing in the resultant control output commands are determined based on a wing angle parameter, wherein the wing angle parameter corresponds to a collective angle of the first and second wing members and represents the state of the transition between the rotating rotorcraft flight mode and the non-rotating rotorcraft flight mode.

11. The hybrid aerial vehicle according to claim 10, wherein, for a transition from the rotating rotorcraft flight mode to the non-rotating rotorcraft flight mode, the amount of the first control output commands and the amount of the second control output commands for contributing in the resultant control output commands are determined based on a first sigmoid function configured as a function of the wing angle parameter, and for a transition from the non-rotating rotorcraft flight mode to the rotating rotorcraft flight mode, the amount of the first control output commands and the amount of the second control output commands for contributing in the resultant control output commands are determined based on a second sigmoid function configured as a function of the wing angle parameter.

12. The hybrid aerial vehicle according to claim 11, wherein the first sigmoid function is parameterized based on one or more first function parameters comprising one or more of a transition midpoint parameter and a transition width parameter for the first sigmoid function, and the second sigmoid function is parameterized based on one or more second function parameters comprising one or more of a transition midpoint parameter and a transition width parameter for the second sigmoid function.

13. The hybrid aerial vehicle according to claim 12, wherein the one or more first function parameters and the one or more second function parameters are optimized based on an objective function.

14. The hybrid aerial vehicle according to any one of claims 9 to 13, wherein the second flight mode controller is configured to output the second control output commands further for controlling a flight of the hybrid aerial vehicle for the fixed-wing flight mode, and the mode transition controller is further configured to, for a transition between the nonrotating rotorcraft flight mode and the fixed-wing flight mode, control an amount of target base pitch angle modulation based a state of the transition between the non-rotating rotorcraft flight mode and the fixed-wing flight mode.

15. The hybrid aerial vehicle according to claim 14, wherein, for a transition from the non-rotating rotorcraft flight mode to the fixed-wing flight mode, the amount of target base pitch angle modulation is determined based on a third sigmoid function configured as a function of the state of the transition between the non-rotating rotorcraft flight mode and the fixed-wing flight mode, and for a transition from the fixed-wing flight mode to the non-rotating rotorcraft flight mode, the amount of target base pitch angle modulation is determined based on a fourth sigmoid function configured as a function of the state of the transition between the non-rotating rotorcraft flight mode and the fixed-wing flight mode.

16. The hybrid aerial vehicle according to claim 15, whereinthe third sigmoid function is parameterized based on one or more third function parameters comprising one or more of a transition midpoint parameter and a transition width parameter for the third sigmoid function, and the fourth sigmoid function is parameterized based on one or more fourth function parameters comprising one or more of a transition midpoint parameter and a transition width parameter for the fourth sigmoid function.

17. The hybrid aerial vehicle according to claim 16, when dependent on claim 13, wherein the one or more first function parameters, the one or more second function parameters, the one or more third function parameters and the one or more fourth function parameters are optimized based on the objective function.

18. The hybrid aerial vehicle according to any one of claims 14 to 17, wherein the first control output commands comprise a first flap actuator output command for controlling the first flap actuator, a second flap actuator output command for controlling the second flap actuator, a first thrust output command for controlling the first thrust unit and a second thrust output command for controlling the second thrust unit, and the first flight mode controller is configured to generate the first and second flap actuator output commands of the first control output commands for controlling the flight of the hybrid aerial vehicle for the rotating rotorcraft flight mode based on a pitch angle and a roll angle of a desired tip path plane of the hybrid aerial vehicle.

19. The hybrid aerial vehicle according to any one of claims 14 to 18, wherein the second control output commands comprise a first flap actuator output command for controlling the first flap actuator, a second flap actuator output command for controlling the second flap actuator, a first thrust output command for controlling the first thrust unit and a second thrust output command for controlling the second thrust unit, and the second flight mode controller is configured to generate the second control output commands for controlling the flight of the hybrid aerial vehicle for the non-rotating rotorcraft flight mode and for the fixed-wing flight mode based on a desired quaternion attitude of the hybrid aerial vehicle.

20. The hybrid aerial vehicle according to claim 19, wherein the second flight mode controller is configured to: generate a virtual torque command based on the desired quaternion attitude of the hybrid aerial vehicle; and generate the second control output commands based on the virtual torque command and a control allocation matrix.

21. A method of operating the hybrid aerial vehicle according to claim 9, the method comprising: controlling, by the wing controller and in response to the transition control signal being for transitioning the hybrid aerial vehicle from the rotating rotorcraft flight mode to the nonrotating rotorcraft flight mode, the wing actuator to rotate the first and second wing members about the lateral axis of the central member from the extended state to the retracted state, and controlling, by the wing controller and in response to the transition control signal being for transitioning the hybrid aerial vehicle from the non-rotating rotorcraft flight mode to the rotating rotorcraft flight mode, the wing actuator to rotate the first and second wing members about the lateral axis of the central member from the retracted state to the extended state.

22. The method according to claim 21, further comprising: outputting, by the first flight mode controller, the first control output commands for controlling the flight of the hybrid aerial vehicle for the rotating rotorcraft flight mode; outputting, by the second flight mode controller, the second control output commands for controlling the flight of the hybrid aerial vehicle for the non-rotating rotorcraft flight mode; and controlling, by the mode transition controller and for the transition between the rotating rotorcraft flight mode and the non-rotating rotorcraft flight mode, the amount of the first control output commands and the amount of the second control output commands for contributing in the resultant control output commands for controlling the flight of the hybrid aerial vehicle, based the state of the transition between the rotating rotorcraft flight mode and the non-rotating rotorcraft flight mode.

23. The method according to claim 22, wherein said controlling, by the mode transition controller and for the transition between the rotating rotorcraft flight mode and the non-rotatingrotorcraft flight mode, the amount of the first control output commands and the amount of the second control output commands for contributing in the resultant control output commands comprises: determining, for the transition between the rotating rotorcraft flight mode and the nonrotating rotorcraft flight mode, the amount of the first control output commands and the amount of the second control output commands for contributing in the resultant control output commands based on a wing angle parameter, wherein the wing angle parameter corresponds to a collective angle of the first and second wing members and represents the state of the transition between the rotating rotorcraft flight mode and the non-rotating rotorcraft flight mode.

24. The method according to claim 23, wherein said determining, for the transition between the rotating rotorcraft flight mode and the non-rotating rotorcraft flight mode, the amount of the first control output commands and the amount of the second control output commands for contributing in the resultant control output commands comprises: determining, for a transition from the rotating rotorcraft flight mode to the non-rotating rotorcraft flight mode, the amount of the first control output commands and the amount of the second control output commands for contributing in the resultant control output commands based on a first sigmoid function configured as a function of the wing angle parameter; and determining, for a transition from the non-rotating rotorcraft flight mode to the rotating rotorcraft flight mode, the amount of the first control output commands and the amount of the second control output commands for contributing in the resultant control output commands based on a second sigmoid function configured as a function of the wing angle parameter.

25. The method according to claim 24, wherein the first sigmoid function is parameterized based on one or more first function parameters comprising one or more of a transition midpoint parameter and a transition width parameter for the first sigmoid function, and the second sigmoid function is parameterized based on one or more second function parameters comprising one or more of a transition midpoint parameter and a transition width parameter for the second sigmoid function.

26. The method according to claim 25, wherein the one or more first function parameters and the one or more second function parameters are optimized based on an objective function.

27. The method according to any one of claims 22 to 26, further comprising: outputting, by the second flight mode controller, the second control output commands for controlling a flight of the hybrid aerial vehicle for the fixed-wing flight mode, and controlling, by the mode transition controller and for a transition between the nonrotating rotorcraft flight mode and the fixed-wing flight mode, an amount of target base pitch angle modulation based a state of the transition between the non-rotating rotorcraft flight mode and the fixed-wing flight mode.

28. The method according to claim 27, wherein said controlling, by the mode transition controller and for the transition between the non-rotating rotorcraft flight mode and the fixed- wing flight mode, the amount of target base pitch angle modulation comprises: determining, for a transition from the non-rotating rotorcraft flight mode to the fixed- wing flight mode, the amount of target base pitch angle modulation based on a third sigmoid function configured as a function of the state of the transition between the non-rotating rotorcraft flight mode and the fixed-wing flight mode; and determining, for a transition from the fixed-wing flight mode to the non-rotating rotorcraft flight mode, the amount of target base pitch angle modulation based on a fourth sigmoid function configured as a function of the state of the transition between the non-rotating rotorcraft flight mode and the fixed-wing flight mode.

29. The method according to claim 28, wherein the third sigmoid function is parameterized based on one or more third function parameters comprising one or more of a transition midpoint parameter and a transition width parameter for the third sigmoid function, and the fourth sigmoid function is parameterized based on one or more fourth function parameters comprising one or more of a transition midpoint parameter and a transition width parameter for the fourth sigmoid function.

30. The method according to claim 29, when dependent on claim 26, wherein the one or more first function parameters, the one or more second function parameters, the one or more third function parameters and the one or more fourth function parameters are optimized based on the objective function.

31. The method according to any one of claims 27 to 30, wherein the first control output commands comprise a first flap actuator output command for controlling the first flap actuator, a second flap actuator output command for controlling the second flap actuator, a first thrust output command for controlling the first thrust unit and a second thrust output command for controlling the second thrust unit, and the method further comprises generating, by the first flight mode controller, the first and second flap actuator output commands of the first control output commands for controlling the flight of the hybrid aerial vehicle for the rotating rotorcraft flight mode based on a pitch angle and a roll angle of a desired tip path plane of the hybrid aerial vehicle.

32. The method according to any one of claims 27 to 31, wherein the second control output commands comprise a first flap actuator output command for controlling the first flap actuator, a second flap actuator output command for controlling the second flap actuator, a first thrust output command for controlling the first thrust unit and a second thrust output command for controlling the second thrust unit, and the method further comprises generating, by the second flight mode controller, the second control output commands for controlling the flight of the hybrid aerial vehicle for the non-rotating rotorcraft flight mode and for the fixed-wing flight mode based on a desired quaternion attitude of the hybrid aerial vehicle.

33. The method according to claim 32, further comprising: generating, by the second flight controller, a virtual torque command based on the desired quaternion attitude of the hybrid aerial vehicle, wherein said generating, by the second flight controller, the second control output commands is based on the virtual torque command and a control allocation matrix.

34. A method of forming the hybrid aerial vehicle according to any one of claims 1 to 20.

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