Patent ID: 12240595

DETAILED DESCRIPTION OF THE INVENTION

The present disclosure discusses inventive aspects that yield improvements in the carbon footprint, flight endurance time, and precision of control of velocity and position for rotorcraft. The principles discussed herein can apply to both crewed or uncrewed vertical takeoff and landing flight vehicles. As will be discussed in more detail below in connection with specific embodiments, a rotorcraft is contemplated having a non-rotating fuselage portion, a rotor attachment ring, and a plurality of rotating rotors extending radially outwardly from the rotor attachment ring. The term rotor in the present disclosure refers to the structure that is also sometimes referred to in the art as a rotary wing or rotor blade. Each rotor may be independently capable of storing electrical energy for flight, harvesting solar energy for flight, providing motive power for flight, and providing the specific aerodynamic surface control and propulsion control so that the velocity and position of the aircraft as a whole are controllable and useful.

With initial reference toFIGS.1-4, an embodiment of an uncrewed utility air vehicle100comprises a minimalistic cylindrical fuselage102that may be configured to house flight control, communications, and mission payload components440, as well as batteries that store energy necessary to power the various components. An array of solar photovoltaic cells104may be arranged on the upper surface of fuselage102to harvest any available solar energy. A rotor attachment ring106may be rotatably attached to the fuselage102and fixedly attached to the depicted individual rotors108a,108b,108c. The rotors108a-cmay employ semi-symmetrical airfoil sections and may each have an array of solar photovoltaic cells120arranged on the upper surface. Mounted at a location along the outer portion or at the distal tips of each rotor, there may be provided one or more aerodynamic fences or “winglets”110a,110b,110cconfigured to increase the apparent aspect ratio, and thus reduce the induced drag of the rotor. These winglets may be arrayed above, below, or both above and below the main corpus of each of the rotors108a,108b,108cdepending on the particular embodiment. A plurality of winglets110a,110b,110cmay also be arrayed at the distal tips of rotors108a,108b,108c. As will be discussed below in connection with additional embodiments, by articulation of certain parts, the winglets110may be caused to generate a variety of vertical, radial, and tangential forces. An improvement in precision of control is realized using this method because a maneuvering or corrective force may be generated almost instantaneously instead of a force created by tipping the entirety of the rotor disk, which suffers a delayed and out-of-axis response due to rotational inertia and gyroscopic precession effects.

With particular reference toFIGS.1and2, a thrust generator112a,112b,112ccan be mounted either within or external to an outer section of each of the rotors108a,108b,108c. In preferred embodiments, and as will be discussed in more detail below, the thrust generators112can be variably throttled in different sectors of a particular rotor's rotation about a spin axis118of the utility air vehicle100to maintain the rotational speed of the rotor assembly as well as to generate direct, net lateral force on the utility air vehicle100. Like the direct forces generated by the control surfaces on the winglets110mentioned above, variable throttling of the thrust generators112accrues the control advantage of near-instantaneous generation of lateral control forces when needed.

In the embodiment illustrated inFIG.1, the thrust generator112comprises an electric motor turning a propeller to produce thrust. The thrust generator112is attached to the leading edge of the rotor108. However, it is to be understood that the thrust generator may be attached to the leading edge of the rotor, the trailing edge of the rotor, or above, below, or within the corpus of the rotor blade structure with any accommodation for entraining and accelerating airflow to produce thrust. As such, it is contemplated in the embodiment illustrated inFIG.1that the motor and propeller can act in either the tractor or pusher configuration.

Each winglet110a,110b,110ccan include one or more control surfaces114a,114b,114c, each of which can generate a drag force, a lateral ‘lift’ force, or a combination of both as its attached rotor108a,108b,108cspins through different sectors around spin axis118as needed for control of air vehicle100. Further, one or more lift-force control surfaces116a,116b,116ccan be included with the rotor108, and can be configured to produce both collective and cyclic lift force through rotors108a,108b,108c. In the embodiment illustrated inFIGS.1and2, the lift-force control surfaces116a,116b,116care connected to the trailing edge of the rotors108a,108b,108cand may operate much the same as an aileron or flap control surface on a conventional airplane wing. It is to be understood that each lift-force control surface116a,116b,116cserves to adjust the upward lift force of the rotors108a,108b,108cas they spin through different sectors of their rotation about fuselage102to facilitate control of velocity and position of air vehicle100. The lift-force control surfaces may also be arrayed in a multitude of locations and be of additional configurations including but not limited to leading-edge control surfaces, trim- or servo-tab surfaces, camber changing facilities, or thrust deflectors placed in the wake of the thrust generators112a,112b,112cto change the circulation of airflow around the rotor blade and therefore change the lift generated. It is contemplated that various embodiments of each of these control elements, including multiple instances distributed along the span or chord of the rotor, will likely have beneficial or detrimental effects on different aspects of flight and performance.

It is contemplated that a rotorcraft air vehicle100in accordance with this disclosure can be configured in a broad range of sizes and diameters. In a preferred embodiment the air vehicle100has a rotor disk span in the range of approximately 3-8 meters, and more preferably about 4-6 meters. Preferably, the rotors may spin at a speed on the order of 80-150 revolutions per minute (RPM), more preferably about 90-120 RPM, and most preferably about 100 RPM. This is in high contrast with most conventional helicopters which have a rotor speed from about 200 RPM to about 600 RPM, or quad-rotor propellers which have speeds approaching or exceeding 10,000 RPM.

Construction of the air vehicle100may employ materials and methods as are in keeping with best practices currently in use in the industry. It will be recognized by one with skill in the art that the choice of elements and the overall size and configuration of any particular embodiment of the invention may reflect engineering trade-offs made to optimize performance in particular flight missions. Further, as more advanced and advantageous materials, components, and manufacturing techniques become available they may be employed as best suited to embodiments of the present invention. A wide variety of architectures and configurations may be possible without departing from the spirit and scope of the present invention.

With specific reference toFIG.2, the fuselage comprises a non-rotating hub202encircled by a rotor attachment ring106. The fuselage102supports the rotor attachment ring106, to which the rotors108a,108b,108care attached so as to extend radially outwardly from the rotor attachment ring106. In addition to the rotor attachment ring106, the fuselage102supports the rotors108a,108b,108cat all times when the rotors are stationary (e.g. the air vehicle100is on the ground and not in flight) or are not yet developing enough lift to support themselves. This particular simplified embodiment depicted provides for a rotor ring support collar204, securely attached to the surface of the non-rotating hub202on the exterior surface of the fuselage102, which supports the rotor attachment ring106and its attached rotors.

Low-friction bearings or bearing surfaces may be provided to minimize drag torque while the fuselage remains pointing along the same direction or heading, supported by the rotor attachment ring106and its spinning rotors108a,108b,108cwhile in flight. Likewise, the fuselage may support the rotor attachment ring106and its spinning rotors108a,108b,108cusing similar bearings or bearing surfaces when not in flight. Bearings notwithstanding, it may be difficult to completely eliminate the drag force at the bearings or bearing surfaces which gives rise to a torque moment which works to slow the rotation of the rotor attachment ring106and its spinning rotors108a,108b,108cand may cause the fuselage102to begin to spin in the same direction as the rotation of rotor attachment ring106and its attached parts.

FIGS.3a,3b, and3cdepict embodiments of anti-rotation structures configured to stabilize and control the rotation angle of the non-rotating hub202and fuselage102with respect to the rotor attachment ring106and its attached rotors108a,108b,108c. The hub202and fuselage102may be made to remain stationary or controllably rotated by one or more of a multitude of devices and methods, some of which are depicted inFIG.3. The skilled artisan will recognize the requirement to control the heading angle of the fuselage102, the function of each of the examples inFIGS.3a,3b, and3c, and that a large variety of methods may be used to accomplish this function without departing from the spirit and scope of the present invention.

FIG.3ais a section view including the fuselage102, the rotor attachment ring106, the rotor ring support collar204, and a friction drive assembly302. The friction drive302is rigidly mounted and housed within the fuselage102and protrudes through an opening308in the non-rotating hub202. The friction drive comprises a semi-pneumatic tire element304and an electric motor306. The tire element304has a compliant, interference fit so that friction helps develop traction forces as it runs against the inner surface of the rotor attachment ring106. The electric motor306may turn the tire element304at a speed necessary to match the rotational speed of the rotor attachment ring106thus keeping the fuselage102always pointing the same direction while the rotor attachment ring106and its attached rotors108a,108b,108cspin at rotational speeds necessary to sustain flight. Alternatively, the electric motor306may turn the tire element304at a speed slightly faster or slower than the rotational speed of the rotor attachment ring106in order to alter the heading angle of the fuselage102as desired.

FIG.3bis a section view including the fuselage102, the rotor attachment ring106, the rotor ring support collar204, and a gear drive assembly312. The gear drive assembly312is rigidly or compliantly mounted and housed within the fuselage102and protrudes through the opening308in the non-rotating hub202. The gear drive assembly312comprises a pinion gear element314and the electric motor306. The pinion gear element314engages a larger ring gear316embedded into the inner surface of the rotor attachment ring106. The electric motor306may turn the pinion gear element314at a speed necessary to match the rotational speed of the rotor attachment ring106thus keeping the fuselage102always pointing the same direction while the rotor attachment ring106and its attached rotors108a,108b,108ccontinuously spin at rotational speeds necessary to sustain flight. Alternatively, the electric motor306may turn the pinion gear element314at a speed slightly faster or slower than the rotational speed of the rotor attachment ring106in order to alter the heading angle of the fuselage102as desired.

FIG.3cis a section view including the fuselage102, the rotor attachment ring106, the rotor ring support collar204, and an anti-torque rotor assembly322. The anti-torque rotor assembly322is mounted to the fuselage102and protrudes beyond the edge of rotor attachment ring106with enough clearance to avoid the spinning rotors108a,108b,108c, much the same as the tail rotor of a conventional helicopter. The anti-torque rotor comprises a structural arm324, an electric motor326, and a fan or propeller328. Power is supplied to the electric motor and modulated as necessary from control systems within the fuselage102. The electric motor326may turn the propeller328at a speed necessary to create a force at the center of action of the propeller for a torque of equal magnitude and opposite direction from the drag torque from the rotational bearings thus keeping the fuselage102always pointing the same direction while the rotor attachment ring106and its attached rotors108a,108b,108cspin at rotational speeds necessary to sustain flight. Alternatively, the electric motor326may turn the propeller328at a speed slightly faster or slower than that necessary to exactly counter the drag torque from the rotational bearings. In this manner it is possible to alter the heading angle of the fuselage102as desired.

While not in flight, the air vehicle100may be supported on landing gear (not depicted) connected to the fuselage102, or the fuselage102may rest directly on the ground or landing surface, supported by a reinforced lower portion of the fuselage102.

With specific reference next toFIGS.2,4a,4cand4d, a vehicle control system may comprise a plurality of sensors401, a central flight control computer, servo-actuators, electronic speed controls, power management electronics, and a plurality of interface devices to facilitate operation. Most of the systems and devices may be commercially available and may be chosen due to function, cost, complexity, ease of maintenance, and deployment considerations.

A systems control computer420, which may comprise multiple digital and analog computing devices, have non-volatile memory storage, and receive command inputs418from a plurality of possible sources, may be enclosed within the fuselage102, to which payload components440, such as a camera and/or sensor array for making observations, may be attached. In the crewed instance, these command inputs418may be human inputs in the form of button press sequences, knob position selections, and flight controls positions. In the uncrewed case, these inputs may come from radio frequency (RF) transceiver417,415data links with one or more remote control stations405on the ground. A transceiver pair417,415may provide wireless communication between the air vehicle100and a remote control station405. In a fully autonomous case, the command inputs118may come from a high level decision engine with command specificity provided by a sort of operational autopilot.

Control of the air vehicle100via the systems control computer420requires a plurality of sensors401so that a high degree of accuracy regarding the state (position, orientation, velocity) of the vehicle may be derived. Fusion of the sensor data may be accomplished using state estimator software elements that model the amount of error inherent in each sensor and place the appropriate weight on each element of the data so that an optimal estimate of the vehicle state is produced. This derived vehicle state may be compared to a commanded or desired state, and control commands may be produced to move the vehicle to the desired state. One skilled in the art will recognize that this core control capability of the system is of prime importance, but that in this regard, myriad embodiments of the present invention can exist and fulfill the basic function without departing from the spirit and scope of the present invention.

The array of sensors employed by the present invention may include, but are not limited to the following examples:

A temperature sensor402and/or a barometric pressure sensor404may be used to provide atmospheric information to the systems control computer, which may in turn be used to optimize vehicle performance, considering factors such as the air density in which the air vehicle100is flying.

Position, and by derivation of the change in position over time, velocity may be derived using sensors such as an optical flow velocity sensor410and a global navigation satellite system (GNSS) receiver406. The accuracy of the GNSS receiver may be significantly improved using RF-linked signals408from ground stations for facilities such as wide area augmentation service (WAAS) and real-time kinematic (GPS RTK) methods with Radio Technical Commission for Maritime (RTCM) corrections. Distance above the ground may be provided by an acoustic method altimeter (typically ultra-sound)412or an electromagnetic spectrum altimeter (typically laser light or radio frequency)414. Finally, orientation, acceleration, and by integration, velocity information may be provided by a plurality of specific sensors to measure acceleration, gyroscopic angular rates, and strength and orientation of the local magnetic field. Collectively, with the appropriate filtering and fusion of the raw data from these sensors, an integrated module is formed which often referred to as an inertial measurement unit (IMU) or inertial measurement system416.

The systems control computer420, which may be housed within the fuselage102, performs multiple control functions for the safe and effective operation of the air vehicle100. A supplementary power source432for operation of the systems within the fuselage102may be provided by, for example, its own array of photovoltaic cells104for harvesting solar energy. The supplementary electric power may be communicated to a power supply433, which may include a power conditioner such as a charge controller434to convert supplementary electric power to a level appropriate to charge a battery energy storage438carried within the fuselage102. The power supply433may also include a battery management system436configured to monitor and control the battery energy storage438which, it is to be understood, may be made up of one or more modular packs of batteries. A fuselage-side signal coupling442can be configured to communicate, through transmit/receive coupling stations443,445with a rotor-side signal coupling444to facilitate data exchange therebetween.

Chief among the control functions of the systems control computer420are the flight control system functions. At the lowest level, flight control functions termed inner-loop flight control functions serve to augment the controllability and stability of the vehicle. These functions augment any natural stability of the air vehicle100to resist perturbations due to, e.g. local wind gusts, by providing for active changes in the control surfaces and thrust generators of the vehicle to more quickly return to the desired position, velocity, and orientation commanded prior to the disturbance. Inner-loop flight controls may additionally enhance the vehicle's response to control inputs. From the control perspective, above the inner-loop flight control functions, are the outer-loop and operational autopilot flight control functions. These functions may actively plan for a sequence of positions and velocities, and derive the vehicle orientations needed to accomplish the maneuver. These functions may further calculate required flight profiles that are multiple flight sequences required for mission success based on local terrain, environmental conditions, vehicle performance limitations, or regulatory limitations. It is to be understood that various embodiments of these control elements can be employed so long as they provide for effective operation of air flight vehicles.

Control of the vehicle provided by the systems control computer420may be facilitated by elements such as a fuselage heading control angle system422(examples of which were described inFIG.3a,3b,3c). The systems control computer420may additionally manage communication with and state of additional vehicle systems424such as anti-collision lighting, a landing configuration system (gear, lighting, special sensors), payload control, communications links, and energy management. Control of vehicle subsystems by the systems control computer420may be facilitated by elements such as command signals sent to the rotors108a,108b,108c, and may be derived by the systems control computer420in the form of collective and cyclic rotor commands426, which generally apply to all rotors, and individual rotor commands428which may apply specifically to individual rotors based on differences in individual rotor status430. Cyclic rotor commands may take the form of a two-value command which includes 1) a maximum amplitude command and 2) the phase angle or rotor sector in which the maximum control amplitude is to be applied. Elements of individual rotor status430may include such data as a current rotor RPM speed, rotor battery temperature and charge state, and rotor rate of charge or discharge of the rotor battery system.

Communication of the control and status system between the systems control computer420and the individual rotor systems450may be accomplished in a variety of ways. Data must flow to and from the rotating rotor attachment ring106and rotors108a,108b,108c, from and to the systems in the fuselage102which may not be rotating at all. Electrically-coupled slip rings may be employed to allow communication between the two parts of the air vehicle100. Short range radio frequency communication may also be used. Optical communication, which uses light energy to send coded messages across short gaps, may also be used to facilitate data transfer to and from the rotors. If the choice of optical communication is constrained to occur during only specific rotation angles when a particular rotor is in a particular sector of its rotation about the spin axis118, then transmit/receive coupling stations443,445may be arrayed to ensure adequate separation in time between communication between the fuselage and each individual rotor. Further, data communication can be configured so as to be limited to communicating bursts of data when transmit/receive coupling stations443,445are aligned. Facility for power and signal conditioning for transmit/receive coupling stations445,443may be provided in each rotor as rotor-side signal coupling444and in the fuselage102as fuselage-side signal coupling442.

Still further, in some embodiments, data communication between the fuselage102and rotors108may use fundamentally different operating principles than data communications between the fuselage102and a remote control station405. For example, communications between the fuselage102and the remote control station405can be accomplished via radio frequency wireless signals, but communications between the fuselage102and rotors108can be accomplished wirelessly by optical, line of sight communication or by wired communications via slip rings. As such, communication between the fuselage102and rotors108is unlikely to interfere with communication between the fuselage102and remote control station405.

In parallel with control functions, the systems control computer420may provide elements such as safety and logistical function in the operation of air vehicle100. Communication between the systems control computer420and the rotors108a,108b,108cmay be subject to data transmission error and interruption. The communications protocol used between them may incorporate techniques such as data bit parity, guard bits or words, checksum verification, or self-correcting codes to ensure the integrity of the data sent in individual packets. It may employ techniques such as rolling codes, cyclical redundancy codes, and inter-packet timing to ensure that fresh data is always available or stale data is recognized and compensated for. In the event of a sustained loss of data integrity or currency, the control protocols of either or both of the systems control computer420or the individual rotor systems450may call for configuration of the vehicle or parts of the vehicle for safest return to a low-risk and/or secure landing area.

With specific reference toFIGS.4aand4c, electric power for components in the fuselage102may be provided by the power supply433comprising the battery charge controller434, the battery management system436, and the battery energy storage438. A supplementary power source432(which, for example, may comprise an array of solar photovoltaic cells104), may provide electrical power supplementary to the energy stored in the battery energy storage438, and may communicate electrical power to the power supply433. When an excess of electrical power is available from the supplementary power source432, the excess electrical current may flow, in a manner regulated and controlled by the battery charge controller434, into the battery energy storage438for use at a later time. The battery energy storage438may further be protected by the battery management system436, which may protect the battery energy storage438from hazardous conditions such as overcharge, overdischarge, or over-temperature, and help mitigate performance degrading conditions such as excessively low operating temperature or excessively high operating temperature. In some embodiments, supplementary power source432may comprise an array of photovoltaic solar cells configured to harvest solar energy from incident sunlight, as available. In other embodiments, supplementary power source432may comprise a stored energy source and a means to convert the stored energy into useable electrical power. One example, without limiting others is stored hydrogen, the reaction of which in a hydrogen fuel cell, may generate electrical power. Another example, without limiting others is stored petrochemical fuel such as diesel fuel, which may be burned in an internal combustion engine which may be configured to turn an electrical generator to generate electrical power.

With specific reference next toFIGS.4band4d, a power supply453of the individual rotor system450may include power control elements very similar to those in the fuselage as detailed inFIGS.4aand4c. A battery energy storage458may be present, as well as a supplementary power source452(which, for example, may comprise an array of individual solar cells120) which may provide electrical power supplementary to the energy stored in the battery energy storage458. When an excess of electrical power is available from the supplementary power source452the current may flow, in a manner regulated and controlled by a power conditioner such as a battery charge controller454, into the battery energy storage458for use at a later time. The battery energy storage458may further be protected by a battery management system456, which may protect the battery energy storage458from hazardous conditions such as overcharge, over-discharge, or over-temperature, and help mitigate performance-degrading conditions such as excessively low operating temperature or excessively high operating temperature. In some embodiments the battery energy storage458may comprise a plurality of separate battery packs458a,458b,458cdistributed along the length of the associated rotor108so as to distribute battery weight within the length of the rotor108. In some embodiments, the supplementary power source452may comprise an array of photovoltaic solar cells configured to harvest solar energy from incident sunlight, as available. In other embodiments, the supplementary power source may comprise a stored energy source and a means to convert the stored energy into useable electrical power. One example, without limiting others, is stored hydrogen, the reaction of which in a hydrogen fuel cell, may generate electrical power. Another example, without limiting others, is stored petrochemical fuel such as diesel fuel, which may be burned in an internal combustion engine which may be configured to turn an electrical generator to generate electrical power.

Control of the individual rotor system450may be provided by a rotor blade controller460which may comprise multiple digital and analog computing devices. The rotor blade controller460may have facility for non-volatile storage of, for example, program code and calibration data, as well as working memory to be used for calculations and storage of ongoing status data. The rotor blade controller460receives commands and sends status information from/to the systems control computer420located in the fuselage102of the air vehicle100. As disclosed in the discussion forFIG.4a, communication between systems control computer420and any individual rotor450may be accomplished in a variety of ways through rotatably free connections.

Specifically germane to the individual rotor system450is the determination of the rotation phase angle or sector of its rotation about the spin axis118. In operation, the systems control computer420may issue a command for a cyclic maximum control effect with respect to its own orientation. To appropriately apply the cyclic control effect throughout the rotation about the spin axis118requires the individual rotor to determine its rotational position relative to the fuselage102heading orientation. Determination of a particular rotor's rotational position may be accomplished in a variety of ways including, but not limited to, electro-mechanical means such as one or more cams or ramps situated on the non-rotating hub202surface so as to engage one or more limit switches mechanically connected to the rotor attachment ring106and electrically connected to the rotor blade controller460so as to provide an electrically detectable indication of rotational position of the rotor with respect to the fuselage102. Similarly, rotational position detection may be accomplished using contactless means such as optical photo-detectors or magnetic (e.g. Hall-effect) detectors. As disclosed in the discussion ofFIG.4a, cyclic control commands may be received from the systems control computer420as two quantities: a commanded maximum amplitude and a rotational angle at which the maximum commanded amplitude of control is to be applied. The amount of control amplitude applied is the product of the maximum commanded amplitude and the calculated, derived, or specified value of the control profile at any position of the particular rotor as it rotates about the spin axis118.

Conventional helicopters typically employ a swashplate to cause cyclic control changes. The swashplate directly and mechanically modifies cyclic angle of attack of the rotor blades. The profile is effectively that of a cosine function where the maximum commanded positive control deflection occurs at the maximum deflection angle of the swashplate and the maximum negative (or minimum positive) control deflection occurs at a point 180 degrees around the rotation circle from the maximum deflection angle of the swashplate.

In the illustrated embodiment, the linkage between the controllers and rotors is informational, not a direct, mechanical one. As such, the profile of cyclic control effect can have myriad additional variations. These include but are not limited to profiles such as clipped or saturated cosine functions in which the maximum control deflection is sustained through a larger arc about the spin axis118instead of just touching the maximum control deflection at a single point or very small portion of the arc about the spin axis. Likewise, (subject to physical constraints, servo-actuator movement rates, and the like) the cyclic control effect may take the profile of something like a sawtooth or square wave. And because the specification of maximum control effect is in the form of digital information, the actual effect on the various control elements of each individual rotor may additionally be implemented some rotation angle ahead of or behind the desired control effect position to achieve control results perfectly in keeping with the intended command input from the systems control computer420. This ability to modify the shape of the control profile, and to ‘lead’ or ‘lag’ control inputs for desired effect may additionally be implemented in real time by an adaptive software element of either or both the systems control computer420or the rotor blade controller460in response to changing atmospheric conditions or vehicle performance requirements.

Implementing desired control functions may be enabled by elements such as electromechanical servo-actuators462, electronic speed controllers464, or auxiliary devices466in response to control signals provided by the rotor blade controller460. Auxiliary devices466may include, but are not limited to, systems such as anti-collision lighting, aimable communications antennae, visual status indicator lighting, retractable landing gear, and controllable payload features. An auxiliary payload476may be carried in individual rotors. Communication between the rotor blade controller460and the various components may be accomplished using a methods and communications protocols appropriate to the design and environmental conditions. Examples include, but are not limited to, methods and protocols such as pulse-width modulated power, RS-232c, RS-485, I2C, SPI, and CAN BUS.

The servo-actuators462may be located in close proximity to direct-lift or servo-tab control surfaces468(comprising elements such as116and526). There may exist a direct, mechanical connection between the servo-actuator462and the direct-lift or servo-tab control surfaces468and configured to change the location or orientation of the direct-lift or servo-tab control surfaces468in order to effect the desired control function, which will ultimately be used to control the position, velocity, and orientation of the air vehicle100. Likewise, servo-actuators462may be mechanically coupled to direct lateral force control surfaces470(comprising elements such as114,704, and706) to effect desired control over the state of air vehicle100.

Electronic speed controllers464may be employed to convert command signals from systems control computer420through the rotor blade controller460into electrical power signals required to drive one or more electric motor(s)472. The electric motor(s)472may turn fans or propellers474, of either fixed or variably-controlled pitch, to produce aerodynamic forces which will control and sustain rotor ring RPM about spin axis118and produce variable aerodynamic forces at different rotation angles or sectors about spin axis118to give rise to net direct lateral forces to be used in the control of air vehicle100.

The rotor blade controller460may control the auxiliary payload476which may be additionally carried within or attached to the rotor. Properly done, these rotor-carried payloads might range from communications gear and electromagnetic spectrum collection devices (cameras, antennae, magnetic anomaly detectors) to droppable objects intended for delivery at particular locations.

It is to be understood that the physical location and configuration of the various components in the illustrated embodiments are shown by example, and it is contemplated that other specific arrangements can be employed, and may be dependent on the intended flight profiles and missions of the particular embodiment of the air vehicle100. Variations beyond the embodiments illustrated may be made without departing from the spirit and scope of the present disclosure.

With reference next toFIG.5a, in a preferred embodiment the shape of the rotor108may be optimized to best accommodate solar cells120, which may cover the entirety of the practical area of the upper surface of the rotor108. The present embodiment provides no mechanical in-flight pitch adjusting mechanism at the root of the rotor blade, such as the linkages to a swashplate mechanism on a conventional helicopter. Instead, a simple, secure attachment may provide structural, mechanical, and data connection between the rotor attachment ring106and each rotor108, providing significant weight savings. Most preferably, the rotor108is rigidly attached to the rotor attachment ring106at its root. In this example, the rotor108has receptacles502which accept protruding tangs504from the rotor attachment ring106. The coupling of the tangs504into the receptacles502may be additionally secured by mechanical fasteners. In other embodiments the protruding tang section may be on the rotor108and the receptacle may be on the rotor attachment ring106. It is contemplated that other structure can also be employed so as to detachably attach each rotor108to the rotor attachment ring106.

Continuing with reference toFIG.5a, the rotor108is elongated, and includes an inboard section508and an outboard section512. The rotor inboard section508in this embodiment has a single angle of incidence510depicted to provide for simplicity of construction. In other embodiments the rotor108may have a continuously variable angle of incidence for optimum angle of attack at the design condition (primarily rotor RPM at a particular all-up weight of the vehicle100) for each spanwise portion of the rotor. The inboard section508of the rotor may be separated from the outboard section512of the rotor by a winglet, or an aerodynamic fence506, which is provided to limit spanwise airflow and limit the formation of aerodynamic vortices which can induce additional aerodynamic drag on the rotor108.

The outboard section512of the rotor may have a reduced, constant angle of incidence514with respect to that shown by510owing to the increased relative wind velocity experienced by the outboard section512, which causes a lower required angle of attack for optimum flight conditions. As with the inboard section508, the outboard section512may have a continuously variable angle of incidence constructed into the structure for optimum flight performance at the design flight condition. Depicted inFIG.5ais a rotor outboard section512with a lift-force control surface116affixed at its trailing edge. This lift-force control surface116may be adjusted by electromechanical servo-actuators mounted within or on the corpus of the rotor108to an average deflection amount to control the total lift generated by the rotors, and it may be further adjusted to different positions by servo-actuators as the rotor spins through different sectors around spin axis118to produce maneuvering forces and moments generated by the rotors.

Located at a point along the rotor outboard section512is the thrust generator112depicted in this embodiment as an electric motor and propeller. In one preferred embodiment the thrust generator112may be oriented to produce thrust along a line540both perpendicular to a line extending radially outward from the center of rotation542and perpendicular to the spin axis118of the rotor system of the vehicle100. This arrangement maximizes the amount of thrust available to maintain rotor RPM.

With reference next toFIG.5b, in another embodiment, the rotor outboard section512comprises a feathering wing section. For clarity, this example does not include a winglet at the distal tip of the rotor. Various embodiments may include or not include the winglet without departing from the spirit or scope of the present disclosure.

An outer feathering wing section520may be free to rotate about a feathering axis522and may utilize a servo-tab or full-flying aerodynamic stabilizer assembly524to determine the rotational position of wing section520about the feathering axis522. It is to be understood that the stabilizer assembly acts much the same as a trim-tab on aerodynamic or hydrodynamic surfaces. The stabilizer assembly524is mounted away from the main body of wing section520so as to be in relatively undisturbed airflow, and mounted aft of the feathering axis522so as to generate a torque about feathering axis522. In this example embodiment, the thrust generator112comprises an electric motor driving a propeller arrayed in the pusher configuration.

The stabilizer assembly524may employ an electromechanical servo-actuator to drive a full-flying stabilizer526to rotate about its aerodynamic center527through different angles of attack so that it generates a lift force528, which causes a torque moment about the feathering axis522. This torque moment in turn rotates outboard feathering wing section520to an angle of attack with respect to the relative wind necessary to cause wing section520to generate a desired amount of lift force at any portion of its rotation about the spin axis118. An alternative embodiment may omit the aerodynamic stabilizer assembly524altogether and provide for directly driving the angle of the feathering wing section520about feathering axis522with an electromechanical servo-actuator. Other possible embodiments may include multiple feathering wing sections present in the rotor outboard section512, or a combination of rotor sections distributed spanwise that are a combination of rotor sections which use either, both, or multiple instances of the lift force control surface (ReferenceFIG.5a-116) and/or the full feathering wing section520arrangement. Any section of the rotor that articulates by feathering may better facilitate auto-rotation if power should be lost. It does so by allowing an angle of attack of the outboard wing section520in descent that generates a lift force generally opposite the descent direction while maintaining or increasing rotor RPM.

With next reference toFIG.5c, in another embodiment, a flapping hinge axis530may be provided for at an intermediate location along the span of the rotor108between the inboard section508and the rotor outboard section512. The hinge mechanism may comprise a mechanical hinge mechanism with the conventional parts of leafs, knuckle, and pin. In other embodiments, the hinge mechanism may simply comprise one or more flat, spring-steel straps securely bonded into the structures of rotor inboard section508and rotor outboard section512and configured so as to allow flexure about hinge axis530while substantially constraining movement in other directions and about other axes. As inFIG.5a, this embodiment depicts the thrust generator112in the tractor (pulling) configuration. As inFIG.5b, for clarity, this example does not depict a winglet at the distal tip of the rotor. The illustrated flapping hinge axis530is generally horizontal and perpendicular to the longest axis of the blade of rotor108. While not located at the root of the rotor108, which is the case on a conventional helicopter, the illustrated flapping hinge axis530serves a similar purpose of stability augmentation as the root-located flapping hinge of a conventional helicopter. As the rotor108encounters a faster relative wind due to the free-stream wind speed caused by motion of vehicle100with respect to the surrounding air mass, the rotor108will generate relatively more lift force. This portion of the rotor disk is referred to as the ‘advancing’ side. The flapping hinge with axis530shown allows the rotor outboard section512to rotate upward due to the increased lift force. Centrifugal force experienced by the rotor outboard section512limits the upward flapping angle532. As rotor outboard section512rotates upward it effectively experiences a reduced angle of attack into the relative wind. The reduced angle of attack causes the outboard rotor section to produce slightly less lift. Eventually, and typically at 90 degrees rotation about the spin axis118from the peak increased relative wind direction, the rotor outboard section512ceases to rotate upward about the flapping hinge with axis530and begins to rotate downward, causing an apparent increase in its angle of attack just as its relative wind speed is decreasing due to the free-stream wind speed occurring in a direction opposite the relative wind speed caused by the rotor's rotation about the spin axis118. Thus the lift force generated on the so-called ‘retreating’ side of the rotor disk is caused to be greater than without the flapping hinge. Differences in lift generated on the advancing side of the rotor disk compared to the retreating side can cause significant rolling moments (pitching moments if the free-stream wind speed is from the side) which can upset the controlled flight of the vehicle. So the addition of a flapping hinge causes a decrease in the excess lift otherwise generated on the advancing side of the rotor disk and an increase in the otherwise deficit of lift otherwise generated on the retreating side of the rotor disk to mitigate any rolling (or pitching) moments that might introduce vehicle control issues. Additionally, as the flapping hinge with axis530allows the rotor outboard section512to rotate upward it causes the lift vector534generated by the rotor outboard section512to tilt opposite the direction of relative movement of vehicle100through the surrounding air mass, producing the stabilizing force536, thus reducing the velocity of vehicle100through the surrounding air mass, thus reducing the upsetting forces on the vehicle.

With reference toFIG.6, the thrust generators112a,112b,112cact to maintain the spin RPM of the rotor system while also generating an aggregate net direct lateral force by modulating thrust within different sectors of each rotor's rotation about the spin axis118. Production of direct lateral force is important because it allows for precise, immediate control of the position and velocity of the air vehicle100without requiring tipping of the rotor disk described by the sweep of the rotors in their rotation about the spin axis118. The rotor disk, owing to the spinning mass of the rotors, behaves much like a gyroscope with its attendant resistance to change in orientation and non-intuitive axis of response (gyroscopic precession) to application of orientation-changing forces. These tendencies, while unavoidable for the general modes of flight of vehicles of this sort, can cause significant difficulty when attempting to perform small, precise maneuvers, where immediate, predictable response is required for acceptable control.

As detailed inFIG.4b, the rotor blade controller460may send command signals to the electronic speed controller464, which controls electrical power to the electric motor472, which drives a fixed- or variable-pitch propeller. The signal from the rotor blade controller460for that particular rotor's thrust generator controls the amount of thrust produced for the particular rotor. This signal from the rotor blade controller460may be varied many times throughout the rotor's movement in its rotation about the spin axis118. These variations in thrust serve to stabilize and control the position, velocity, and orientation of the air vehicle100.

InFIG.6, rotors are depicted in a variety of locations as they rotate about the spin axis118of the vehicle. If the desired effect is to create a net force on the air vehicle100that is directed along the net maneuvering force vector602, then command signals, and therefore thrust produced, may be varied to produce an appropriate net thrust force on the vehicle100while still maintaining rotor RPM, which is needed to sustain flight. It is to be understood that the six discrete positions shown are merely examples, and that the rotor blade controller460will issue commands for different thrust amounts on a substantially continual basis as is appropriate for any angular position as the rotor blade108rotates about spin axis118.

At location600athe thrust generator is pointed in a direction where its effect will have the maximum contribution to the desired net maneuvering force vector602so its thrust604amay be commanded to be substantially greater than an average thrust amount. Here ‘average thrust amount’ means that needed to generally sustain or control the rotor RPM. At location600bthe thrust generator is pointed in a direction nearly perpendicular, but slightly opposed to the desired net force vector602so its thrust amount604bmay be somewhat less than the average thrust amount.

Location600cdepicts an angular position with direction of thrust604cnearly perpendicular to the desired net force vector602and slightly opposed to the desired net force vector602. Here the expected amount of thrust production may be slightly less than the average thrust amount.

Locations600d,600e,600frepresent opposite sides of each of the respective angular positions600a,600b,600c. In these positions the respective amounts of thrust604d,604e,604fmay be expected to be roughly the sum of the average thrust amount minus the surplus or deficit thrust amount of their counterpart removed by 180 degrees of rotation across the rotor disk. The thrust604dmay be significantly less than the average thrust amount owing to its direction opposite the desired net force vector602. In fact, the net thrust at location600d, considering the aerodynamic drag of the thrust generator, may actually be opposite (a net drag force) the usual acting direction of its thrust generator.

The thrust generated as604emay be slightly more than the average thrust amount owing to the fact that the direction of thrust generation begins to align with the direction of desired net maneuvering force vector602. Finally, because the direction of thrust604fis somewhat aligned with the direction of desired net force vector602the amount of thrust, similar to at location600e, may be slightly more than the average thrust amount needed to sustain or change rotor RPM.

With reference next toFIGS.7a-e, an embodiment of a rotor outboard section700includes a rotor-tip winglet702having a pair of independently-operable direct force control surfaces704a,704bat its trailing edge. It is to be understood that locating the winglet at the distal tip of the rotor accrues the advantage of reducing induced drag on the rotor, but is not strictly necessary for generating the cyclic forces comprising a net lateral control force described hereafter. Winglets and aerodynamic fences may be arrayed on any portion of the outboard section of a rotor and still generate aerodynamic forces which are useful for control and maneuvering of the aircraft. Positioning the control surfaces704a,704bin various configurations can generate net direct lateral force by modulating lift and drag force at the winglet location within different sectors of each rotor's rotation about the spin axis118. InFIG.7a, the control surfaces704a,704bare in contact with one another and in line with the chord line of the winglet. This represents the minimum drag configuration, which also generates no lateral lift force.FIG.7bdepicts a configuration in which control surfaces704a,704bare deployed symmetrically in their maximum drag configuration, which generates substantially no lateral lift force.FIGS.7cand7ddepict control surfaces704a,704bin configurations where a radially outward lateral lift force and a radially inward lateral lift force are respectively generated.FIG.7eillustrates a configuration of control surfaces704a,704bmade to generate both a significant radially inward lift force and a significant tangential drag force at the same time.

FIG.7fdepicts another embodiment in which a rotor outboard section700features two winglets706a,706barrayed in a ‘Y’ configuration at the distal tip of the rotor outboard section700. Each winglet706a,706bis separately controlled and operated. The body of each winglet706a,706bmay be rotated through angles714a,714brespectively to produce different amounts of ‘lift’ force and ‘drag’ force, which combine to yield a composite force that can be beneficial for vehicle control. Here ‘lift’ force means that force direction perpendicular to the relative wind and ‘drag’ force means that force direction parallel and in the same direction as the relative wind. The bodies of winglets706a,706bmay be rotated differentially or together about axes708a,708b, or independently without direct consideration of the position of the other to generate vertical and radial forces on the rotor outboard section700and, in turn, on the air vehicle. It is to be understood that this configuration works somewhat similarly to the operation of the empennage of a V-tail configured aircraft.

FIG.7gillustrates operation of an embodiment having rotors with the winglet702and control surfaces704a,704bof the configuration discussed in connection withFIGS.7athrough7e. As shown, the control surfaces704a,704bcan be independently manipulated and positioned about the entirety of the rotor arc to generate a desired control effect. In the illustrated example, the desired effect is to create a net force on the air vehicle100that is directed along the net maneuvering force vector710. Command signals for the lateral force control surfaces may be varied to accomplish this, and it is to be understood that any additional drag created by deflection of direct force control surfaces704a,704bmay need to be compensated for by the thrust generators in order to maintain the desired rotor RPM. As detailed inFIG.6, additional thrust may be applied selectively throughout the rotation of the rotors to both maintain rotor RPM and to generate additional net force in the desired direction.

InFIG.7g, location720f, the direct force control surfaces704a,704bare deflected differentially inboard and outboard as shown to produce the composite force722f. The composite force722fresults from the vector combination of the drag force726and the radially outward ‘lift’ force728. In most angular positions, proper direct force control surface deflections will produce ‘lift’ and ‘drag’ forces that will yield a composite force aligned along the direction of the desired net maneuvering force vector710. It is to be understood that the six discrete positions shown are merely examples, and that the rotor blade controller460will issue commands for control surfaces704a,704bdeflection on a substantially continual basis as is appropriate for any angular position as the rotor blade108rotates about spin axis118.

At location720athe direct force control surfaces704a,704bare deflected inboard and together to produce a radially outward lift force722a. In this case the drag force generated is a small fraction of the lift force and has minimal influence on the resultant force722a.

At location720bthe direct force control surfaces704a,704bare deflected differentially—one inboard and the other slightly more outboard so that the combined force722bis substantially along the direction of the desired net maneuvering force vector710.

At location720cthe direct force control surfaces704a,704bare deflected together and radially outward to produce a radially inward force722c. In this case, there is no possible configuration for the control force surfaces that will produce a force that is well-aligned with the desired net maneuvering force vector710.

At location720dthe direct force control surfaces704a,704bare deflected radially outward to produce a force722dthat is directed radially inward. This is the control position in direct opposition of the force vector710on the rotor arc, yet by properly configuring the direct force control surfaces it is possible to produce a force in the direction of the desired net maneuvering force vector710.

At location720ethe rotor is moving nearly in the direction of the desired net maneuvering force vector710, so very little control input is possible that will produce a force in the desired direction. This position can however be used to produce a small force722eto offset any residual forces generated in other parts of the rotor sweep that were not perfectly aligned with the desired net maneuvering force vector710.

FIG.8depicts one example of a long-endurance embodiment of the present invention. The rotor span is approximately 8 meters. The long-endurance air vehicle800may share similarities with utility air vehicle100where economies of scale may be realized using common elements. The surface area and shape of the rotors808a,808b,808cmay be designed to optimize the structural integrity of the rotor, aerodynamic efficiency of the rotor, propulsive efficiency of the rotor, the coverage area of commercially available photovoltaic cells, and use of readily available, mass-produced components. The structure of any embodiment of the present invention may include structural features designed to optimize weight and manufacturing cost by finding appropriate trade-offs between competing factors. One example of this, without limiting others, is the practice of distributing along the span the relatively heavy energy storage batteries required for long endurance to reduce in-flight bending loads on the wing, while mitigating and accommodating the centrifugal force which the mass of the batteries will impose on the rotor structure as it rotates about the spin axis118. Another example, without limiting others, is the choice of rotor blade chord and span length of the various sections to provide rotor blade upper surface area that optimizes lift produced against ready accommodation of a practical number of photovoltaic cells for solar energy collection as a supplementary power source.

FIG.9depicts one possible embodiment of the present invention which may serve as a crewed transportation vehicle900. The rotor span for the example embodiment shown in as vehicle900is approximately 14 meters. In this example embodiment the minimalistic cylindrical fuselage of utility air vehicle100is replaced with one better configured for crew control, comfort, and safety accommodations. Because of the reversibility of electric motors it may be possible to (inefficiently) generate reverse thrust from the thrust generators to reduce the speed of and to quickly stop the spinning of the main rotors to expedite loading and unloading the vehicle.

It is to be understood that the principles discussed herein can be applied to many flight missions and profiles currently performed by other uncrewed and crewed flight vehicles. Carriage of a wide variety of mission payloads (e.g. communications repeaters; intelligence, surveillance, and reconnaissance equipment; critical items to be delivered for humanitarian purposes) and flight profiles (e.g. extreme high-altitude flight to avoid weather, nap-of-the-earth flight to avoid winds or detection by hostile actors) may entail structural adjustments while still applying inventive principles.

The embodiments discussed above have disclosed structures with substantial specificity. This has provided a good context for disclosing and discussing inventive subject matter. However, it is to be understood that other embodiments may employ different specific structural shapes and interactions.

Although inventive subject matter has been disclosed in the context of certain preferred or illustrated embodiments and examples, it will be understood by those skilled in the art that the inventive subject matter extends beyond the specifically disclosed embodiments to other alternative embodiments and/or uses of the invention and obvious modifications and equivalents thereof. In addition, while a number of variations of the disclosed embodiments have been shown and described in detail, other modifications, which are within the scope of the inventive subject matter, will be readily apparent to those of skill in the art based upon this disclosure. It is also contemplated that various combinations or subcombinations of the specific features and aspects of the disclosed embodiments may be made and still fall within the scope of the inventive subject matter. Accordingly, it should be understood that various features and aspects of the disclosed embodiments can be combined with or substituted for one another in order to form varying modes of the disclosed inventive subject matter. Thus, it is intended that the scope of the inventive subject matter herein disclosed should not be limited by the particular disclosed embodiments described above, but should be determined only by a fair reading of the claims that follow.