Patent Description:
The present disclosure relates generally to manned or unmanned aircraft, and more particularly to aircraft capable of vertical and horizontal flight.

A helicopter is an aircraft in which lift and thrust are supplied by one or more horizontal rotors. The advantages of a helicopter include its ability to hover and to take off and land vertically. However, among other things, a helicopter suffers from its relatively poor operating energy efficiency compared to fixed-wing aircraft.

A gyroplane (also known as a gyrocopter or autogyro) is an aircraft that uses an unpowered rotor in autorotation to develop lift. Autorotation is a rotor state in which the rotor derives from the freestream <NUM>% of the power required to rotate it, and the resulting rotation provides lift. In a gyrocopter, forward thrust is typically provided by an engine-driven propeller. However, like a fixed-wing aircraft, a gyrocopter cannot take off and land vertically.

In the case of a gyroplane, as the aircraft goes down the runway and gathers speed, the overhead rotor's shaft is tilted backwards allowing the wind to blow though the rotor in order to start turning. As the rotor reaches a certain RPM, that very rotor becomes a "virtual wing" for the gyrocopter providing lift. Once it reaches the desired RPM, the gyrocopter is ready for takeoff. As the gyrocopter gains speed in the air the angle of attack of the rotor head (virtual wing) is reduced by moving the shaft forward, increasing speed and reducing drag. A gyrocopter does not have the ability to achieve vertical takeoff because it does not have a variable pitch rotor. The pitch of the blades of a gyrocopter is typically zero. In addition to being able to tilt the shaft forwards and backwards, a gyrocopter can also tilt the shaft/rotor head to starboard and port, providing aileron like maneuverability. Also, for a gyrocopter to fly in a stable manner, it has a unique rotor head assembly that can teeter totter on the shaft allowing the blades freedom of movement as they rotate.

A helicopter has a fixed vertical shaft with swash plates and links to the rotor head, allowing the pilot to modify the pitch of the rotor blades, generating lift. Further, it has a propeller mounter vertically at the end of a boom (or a ducted fan), creating a thrust to counter the torque generated by the motor that drives the rotor head. It is known in the state of the art (<CIT>, <CIT>, <CIT>, <CIT> and <CIT>) that there are rotor assemblies for aircrafts that comprise one or more rotor blades and an upper rotor hub assembly that includes at least one mounting member. That at least one mounting member is known to further include a central section, two mounting brackets (each coupled to each of the rotor blades) and a blade pitch adjustment linkage coupled to the mounting brackets and the central section that allows to automatically adjust the blade pitch of the rotor blades based on a rotational velocity of the rotor blades.

In the case of a helicopter the blades drive the air from the top downwards creating thrust and lift. In the case of a gyrocopter the air flows through the blades upwards spinning them and creating lift.

The systems, methods, and devices of the present technology each have several innovative aspects, no single one of which is solely responsible for its desirable attributes disclosed herein. A rotor assembly according to claim <NUM> is provided. Preferable embodiments are defined by the dependent claims.

The above-mentioned aspects, as well as other features, aspects, and advantages of the present technology will now be described in connection with various implementations, with reference to the accompanying drawings. The illustrated implementations are merely examples and are not intended to be limiting. Throughout the drawings, similar symbols typically identify similar components, unless context dictates otherwise.

Generally described, the present disclosure provides aircraft as well as aircraft systems, components, and control methods providing enhanced flight characteristics relative to existing helicopters, gyroplanes, fixed-wing airplanes, and tiltrotor aircraft. Aircraft disclosed herein may be capable of efficient forward flight, hovering, and/or vertical takeoff and landing (VTOL), as well as transitioning between flight modes in flight. For example, an aircraft in accordance with the present technology may be able to take off vertically from a deployment location, transition to a forward flight mode for generally horizontal flight to a remote location, transition to a vertical flight mode to hover at the remote location for an extended time period, transition to the forward flight mode for generally horizontal flight to a landing location (e.g., the deployment location), and transition again to the vertical flight mode to land at the landing location. Thus, the present disclosure provides aerial vehicles capable of extended data gathering or observation at a relatively distant location beyond the range of a traditional helicopter, without requiring a runway for takeoff and landing.

Some aircraft described herein are further configured to be partially disassembled and may include foldable components to provide a more compact configuration for transportation to or from deployment locations. In one example, a central rotor of the aircraft may include one or more pairs of blades attached to an upper rotor hub assembly. The upper rotor hub assembly may be detachable from a lower rotor hub of the aircraft, and may permit the rotor blades to be folded into a substantially parallel configuration such that the central rotor may be transported in a container approximately the same length as an individual rotor blade, without requiring the rotor blades to be individually separated from the upper rotor hub assembly. When the aircraft is to be deployed again, the central rotor may be conveniently unfolded and attached to the lower rotor hub without requiring additional calibration, alignment, and the like.

<FIG> illustrate an example aircraft <NUM> configured for in-flight transition between vertical and horizontal flight modes. The aircraft <NUM> includes a fuselage <NUM>, a mast <NUM> extending generally upward from the fuselage <NUM>, and an empennage <NUM> disposed aft of the fuselage <NUM>. A tiltable central rotor <NUM> is rotatably coupled at an upper end of the mast <NUM>. Tiltably mounted proprotors <NUM>, 136r are rotatably coupled at opposing ends of a boom <NUM> extending laterally from the mast <NUM>.

The fuselage <NUM> is a body section of the aircraft <NUM> and may include an interior volume sized and shaped to hold a payload. For example, the interior volume of the fuselage <NUM> may be used to contain one or more items being transported by the aircraft <NUM>. The fuselage <NUM> may contain one or more reconnaissance c surveillance devices, such as imaging devices (e.g., a visible light camera, infrared camera, thermal camera, still camera, video camera, synthetic-aperture radar, etc.), listening devices, communications devices, or the like. The fuselage <NUM> may further contain at least some of the control systems for the aircraft <NUM>, such as motor, control surface, or tilt servo controllers, autopilot systems, and the like. If the aircraft <NUM> is configured for operation as a remotely piloted unmanned aerial vehicle (UAV) or drone, the fuselage <NUM> may also include a communication system to receive control commands from a remote pilot. An undercarriage <NUM> may be disposed on a side or bottom portion of the fuselage <NUM> for use during takeoff and landing phases of flight, and may include wheeled landing gear, skids, and/or any other suitable type of undercarriage. The fuselage <NUM> and the undercarriage <NUM> may comprise any suitably rigid or semi-rigid materials such as metal, plastic, carbon fiber, wood, fiberglass, etc..

The mast <NUM> extends generally upward from the fuselage <NUM> and supports the central rotor <NUM> disposed at an upper end of the mast <NUM>. In the example illustrated in <FIG>, the mast also includes a forward tilt such that the mast extends upward at a forward angle from the fuselage <NUM>. As shown in <FIG>, the forward tilt of the mast <NUM> may advantageously allow the rotational axis of the central rotor <NUM> to align with the center of gravity of the fuselage <NUM>. In addition, when the proprotors <NUM>, 136r are tilted upward for vertical/hover flight, the central rotor <NUM>, the longitudinal location of the proprotors <NUM>, 136r, and the center of gravity of the fuselage <NUM> are all aligned with a common center of gravity for improved stability in hover or vertical flight. The mast <NUM> may also serve as a central attachment point for other components of the aircraft <NUM>, including the boom <NUM>, empennage <NUM>, and lower rotor hub <NUM>. The mast <NUM> may further include an interior volume in which one or more aircraft components may be disposed. In some examples, energy storage media such as batteries, hydrogen storage, or the like, may be contained within the mast <NUM>. Placement of batteries or hydrogen storage within the mast <NUM> may advantageously keep such relatively heavy components close to the center of gravity of the aircraft <NUM>, resulting in improved stability.

The central rotor <NUM> is rotatably mounted to a top portion of the shaft <NUM> via a lower rotor hub <NUM>. Rotor blades <NUM> are fixed to an upper rotor hub assembly <NUM> coupled to the lower rotor hub <NUM>. The rotor blades <NUM> may have an airfoil profile configured to enhance the production of lift while the central rotor <NUM> is spinning. Although the rotor <NUM> of the example aircraft of <FIG> has four rotor blades <NUM>, it will be understood that other numbers of rotor blades <NUM> may be included. For example, the rotor <NUM> may have <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, or more rotor blades <NUM>.

The lower rotor hub <NUM> is tiltably mounted at the top of the mast <NUM>, for example, on a central rotor control housing <NUM>. The central rotor control housing <NUM> may include one or more servos configured to provide at least fore and aft tilting of the central rotor <NUM> (e.g., by tilting the lower rotor hub <NUM>). In some examples, the central rotor control housing <NUM> further includes one or more servos configured to provide lateral tilting of the central rotor <NUM>. The central rotor control housing <NUM> can also include a rotor motor configured to turn the central rotor <NUM> during some phases of flight (e.g., during vertical and/or transitional flight modes), as described in greater detail below. In some examples, the rotor motor may be configured to turn the central rotor <NUM> at rotational speeds up to typical gyroplane rotor rotational speeds (e.g., up to approximately <NUM>-<NUM> rpm), which may be slower than typical helicopter rotor rotational speeds (e.g., approximately <NUM>-<NUM> rpm or more). The rotor motor may be rotationally coupled to the central rotor <NUM> by a clutch and/or a one-way bearing such that, during forward flight, the central rotor <NUM> can rotate faster than the rotor motor, and such that the central rotor <NUM> can continue rotating by autorotation when the rotor motor is not turning.

The proprotors <NUM>, 136r are configured to provide lift and/or forward thrust, depending on the tilt of the proprotors <NUM>, 136r. The proprotors <NUM>, 136r are tiltably mounted at opposite ends of the boom <NUM>. In some examples, the boom <NUM> includes distal arms <NUM> and 134r, which are individually pivotable along the lateral axis of the boom, such that each proprotor <NUM>, 136r is independently tiltable by pivoting the distal arms <NUM>, 134r of the boom <NUM>. Servos and/or other actuators disposed within a proprotor tilt control housing <NUM> may control pivoting of the distal arms <NUM>, 134r. As described in greater detail below, each proprotor <NUM>, 136r can be independently powered by left and right proprotor motors and may be operable at different relative speeds for enhanced maneuverability. The speeds of both proprotors <NUM>, 136r may further be adjusted collectively so as to propel the aircraft forward at a range of desired speeds. In some examples, the proprotors <NUM>, 136r may have blades featuring a hybrid shape between a propeller shape and a rotor shape so as to operate efficiently in both forward and vertical/hover flight modes.

The empennage <NUM> is mounted at a rear portion of the aircraft <NUM> and includes a horizontal stabilizer <NUM> and vertical stabilizers <NUM>. A longitudinal tail boom <NUM> may fix the empennage <NUM> to the mast <NUM> or fuselage <NUM>. The horizontal stabilizer <NUM> and vertical stabilizers <NUM> provide stability to the aircraft, primarily during horizontal flight. In some examples, the empennage <NUM> may include one or more control surfaces such as an elevator disposed on the horizontal stabilizer <NUM> and/or rudders disposed on the vertical stabilizers <NUM>. In examples having rudders on the vertical stabilizers <NUM>, the vertical stabilizers <NUM> may be placed at a location within the slipstreams of the proprotors <NUM>, 136r to increase rudder effectiveness at low airspeeds.

With reference to <FIG> and <FIG>, and with continued reference to <FIG>, example flight control systems and methods will now be described. <FIG> schematically illustrates an example tilting range of proprotors <NUM>, 136r with respect to the lateral boom <NUM> of the aircraft <NUM>. <FIG> schematically illustrates example VTOL/hover and forward flight tilt positions of the proprotors. Each of the proprotors <NUM>, 136r may have a range of tilt angles of at least <NUM> degrees, and in some cases up to <NUM> degrees or more. For example, proprotors <NUM>, 136r may be tiltable to a VTOL/hover position or range of positions, in which each proprotor <NUM>, 136r is substantially aligned about a vertical or z-axis to produce an upward lifting force. Proprotors <NUM>, 136r may further be tiltable to a forward flight or horizontal flight position or range of positions, in which each proprotor <NUM>, 136r is substantially aligned about a longitudinal or x-axis to produce forward thrust. Within either the VTOL/hover position or forward flight position, each proprotor <NUM>, 136r may be independently tiltable within a range of approximately <NUM> degrees, <NUM> degrees, <NUM> degrees, <NUM> degrees, or more, such as for maneuvering and/or stability as described below. In some examples, the proprotors <NUM>, 136r can be tilted to any tilt angle between <NUM> degrees below horizontal to <NUM> degrees aft of vertical, for a full tilt angle range of approximately <NUM> degrees.

Vertical flight modes, such as VTOL, hovering, may be achieved with the proprotors <NUM>, 136r in a VTOL/hover position, such as wherein the axis of rotation of each of the proprotors <NUM>, 136r is within approximately <NUM> degrees of vertical. In the VTOL/hover position, the spinning proprotors <NUM>, 136r primarily generate an upward lifting force. In addition, the central rotor <NUM> may also be used during vertical flight modes. The central rotor <NUM> may be maintained in a substantially vertical orientation and turned by the rotor motor at a speed sufficient to provide gyroscopic stabilization to the aircraft <NUM> and produce an additional lifting force in addition to the lift provided by the proprotors <NUM>, 136r. For example, turning the central rotor <NUM> at a relatively low rotational speed (e.g., a typical gyroplane rotor rotational speed) can be sufficient to significantly stabilize the aircraft <NUM>. However, in addition to stabilizing the aircraft <NUM>, the central rotor <NUM> when powered may also create torque effects and gyroscopic precession. In some examples, the proprotors <NUM>, 136r may be differentially controlled, and/or may be configured to rotate in a direction opposite the rotation of the central rotor <NUM>, to counteract the torque effect of the central rotor <NUM>. For example, the torque generated by a clockwise turning central rotor <NUM> may tend to cause the body of the aircraft <NUM> to spin counterclockwise. To counteract this torque effect, the left proprotor <NUM> may be tilted slightly forward of vertical while the right proprotor 136r is tilted slightly aft of vertical, such that the proprotors <NUM>, 136r exert a clockwise torque on the body of the aircraft <NUM> (e.g., a yawing moment) that counteracts the torque effect of the turning central rotor <NUM>. In some examples, a yawing moment may be created by varying the speeds of the proprotors <NUM>, 136r. For example, rotating proprotor <NUM> faster than proprotor 136r creates a yawing moment to the right, and rotating proprotor 136r faster than proprotor <NUM> creates a yawing moment to the left. These methods of generating a yawing moment may be utilized for turning and/or for countering the torque generated when the main rotor is powered.

During vertical flight, such as takeoff, landing, hovering, lateral movement, and/or slow forward or backward flight, differential control of the proprotors <NUM>, 136r may further be used to maneuver the aircraft <NUM>. As described above, differential tilting of the proprotors <NUM>, 136r may be used to produce a yawing moment for maneuverability about the vertical axis. Differential rotation speed of the proprotors <NUM>, 136r may be used to produce a rolling moment for maneuverability about the longitudinal axis. For example, powering the left proprotor <NUM> at a higher rotational speed relative to the right proprotor 136r produces a right or clockwise rolling moment; powering the right proprotor 136r at a higher rotational speed relative to the left proprotor <NUM> produces a left or counterclockwise rolling moment. Simultaneous tilting of the proprotors <NUM>, 136r forward or aft of vertical, and/or tilting the central rotor <NUM> forward or aft, produces a pitching moment for maneuverability about the lateral axis. Rolling or pitching the aircraft <NUM> out of a vertical orientation yields a horizontal component of lift which may be utilized for forward, backward, and/or lateral movement during generally vertical flight.

Horizontal or forward flight, including straight and level flight, climbing, descending, turning, and the like, may be achieved with the proprotors <NUM>, 136r in a forward flight position, such as wherein the axis of rotation of each proprotor <NUM>, 136r is within approximately <NUM> degrees of horizontal. In the forward flight position, the spinning proprotors <NUM>, 136r primarily generate forward thrust substantially parallel to a direction of flight. During forward flight, the central rotor <NUM> may be unpowered and turns in free autorotation. Preferably, the central rotor <NUM> has an upward tilt (e.g. between approximately <NUM> degree and <NUM> degrees or more with the higher side of the central rotor <NUM> oriented toward the direction of flight). In some examples, the central rotor <NUM> is tilted at an angle of approximately <NUM> to <NUM> degrees in forward flight. Thus, in forward flight, the aircraft <NUM> performs substantially as a gyroplane, with the powered proprotors <NUM>, 136r generating thrust and the autorotating central rotor <NUM> providing lift. The empennage <NUM> provides directional stability during forward flight.

In various examples, the empennage <NUM> may or may not include control surfaces. For example, the horizontal stabilizer <NUM> may include one or more elevators configured to provide pitch control, and the vertical stabilizers <NUM> may each include a rudder configured to provide yaw control. In some examples, the empennage <NUM> includes only rudders or only an elevator, and in other examples the empennage <NUM> contains neither rudders nor elevators. Instead, any or all of pitch, yaw, and roll may be controlled by the variable tilt and pitch proprotors. Pitch and/or roll may also be controlled at least in part by tilting of the central rotor <NUM>.

Pitch control in forward flight may be achieved by tilting the central rotor <NUM> forward or aft. Pitch control in forward flight may also be achieved by simultaneously tilting both proprotors <NUM>, 136r higher or lower relative to the longitudinal or x-axis. For example, simultaneous upward tilting of the proprotors <NUM>, 136r can produce a nose-up pitching moment, and simultaneous downward tilting of the proprotors <NUM>, 136r can produce a nose-down forward pitching moment.

Roll control in forward flight may be achieved by tilting the central rotor <NUM> left or right. However, controlling roll by tilting the central rotor <NUM> requires a lateral tilting mechanism for the central rotor <NUM>. In some examples, the lower rotor hub <NUM> and/or the upper rotor hub assembly <NUM> may be simplified by providing only fore and aft tilting, and not lateral tilting. In such examples, aircraft roll can be achieved based on differential tilting of the proprotors <NUM>, 136r. For example, tilting the left proprotor <NUM> slightly upward relative to horizontal and/or tilting the right proprotor 136r slightly downward relative to horizontal produces a right or clockwise rolling moment. Tilting the right proprotor 136r slightly upward relative to horizontal and/or tilting the left proprotor <NUM> slightly downward relative to horizontal produces a left or counterclockwise rolling moment.

Yaw control in forward flight may be achieved by providing differential power to the left and right proprotors <NUM>, 136r. For example, varying the relative speeds of the proprotors <NUM>, 136r such that the left proprotor <NUM> turns at a higher rotational velocity than the right proprotor 136r produces a right yawing moment. Varying the relative speeds of the proprotors <NUM>, 136r such that the right proprotor 136r turns at a higher rotational velocity than the left proprotor <NUM> produces a left yawing moment.

In various examples, turning in forward flight may be achieved by yaw control (e.g., by variable relative proprotor speeds), by roll control only (e.g., by variable relative tilt of the proprotors), or by a combination of yaw control and roll control. In some examples, turning with a combination of yaw and roll control may be desirable in order to maintain coordinated flight without sideslip while turning. For example, a right turn may be performed by simultaneously (or substantially simultaneously) tilting the left proprotor <NUM> upward, increasing the rotational speed of the left proprotor <NUM>, tilting the right proprotor 136r downward, and/or decreasing the rotational speed of the right proprotor 136r.

In addition to vertical and forward flight modes, the aircraft <NUM> is further capable of transitioning between forward and vertical flight modes while in flight. Maneuvers for transitioning from vertical flight to forward flight, and from forward flight to vertical flight, will now be described. Advantageously, the aircraft of the present disclosure are capable of transitioning seamlessly between flight modes without sacrificing stability or controllability during the transition.

The aircraft <NUM> may transition from vertical flight to forward flight at various times during a mission, for example, after a vertical takeoff when entering a cruise portion of flight to a remote location, after a period of hovering at the remote location, etc. The transition from vertical flight to forward flight begins with the aircraft <NUM> configured for vertical flight. In this configuration, the proprotors <NUM>, 136r are in the VTOL/hover position illustrated in <FIG>, and the central rotor <NUM> may be turning under power while substantially parallel to the proprotors <NUM>, 136r. To transition to forward flight, the proprotors <NUM>, 136r simultaneously tilt forward into the forward flight position illustrated in <FIG>. At approximately the same time, the central rotor <NUM> is tilted rearward (e.g., higher toward the front of the aircraft and lower toward the rear of the aircraft). If the central rotor <NUM> was powered during the preceding vertical or hovering flight, the rotor motor may continue to turn the central rotor <NUM> until the aircraft is established in forward flight. If the central rotor <NUM> was not powered during the preceding vertical or hovering flight, the rotor motor may activate during the transition in order to spin up the central rotor <NUM> to an appropriate rotational speed to provide lift for forward flight. Once the aircraft <NUM> is established in forward flight, the rotor motor may be deactivated as the relative airflow against the central rotor <NUM> becomes fast enough to cause autorotation of the central rotor <NUM>.

The aircraft <NUM> may transition from forward flight to vertical flight a various times during a mission, for example, upon arrival at a remote location where the aircraft <NUM> will hover for a period of time, upon arrival at a landing site, etc. The transition from forward flight to vertical flight begins with the aircraft <NUM> configured for horizontal flight. In this configuration, the proprotors <NUM>, 136r are in the forward flight position illustrated in <FIG>, and the central rotor <NUM> is turning by free autorotation as the aircraft travels in forward flight. To transition to vertical flight, the proprotors <NUM>, 136r simultaneously tilt upward into the VTOL/hover position illustrated in <FIG>. The central rotor <NUM> is tilted forward (e.g., to a level orientation substantially parallel to the proprotors <NUM>, 136r). In some examples, the proprotors <NUM>, 136r may be tilted beyond vertical, such as between <NUM> degrees and <NUM> degrees aft of vertical, if desired to slow the forward airspeed of the aircraft <NUM>. As the airspeed decreases and/or as the central rotor <NUM> tilts back to a level orientation, the autorotation of the central rotor <NUM> may decrease and/or stop. The rotor motor may be activated to turn the central rotor <NUM> to provide additional lift and/or stabilization to the aircraft <NUM> during the vertical flight phase.

Throughout the preceding disclosure, proprotors <NUM>, 136r are described as being independently tiltable and controllable in order to achieve various flight control functions. <FIG> further illustrate example components configured to implement the proprotor control features described herein. <FIG> depicts a medial portion of the lateral boom <NUM> of an example aircraft <NUM> including proprotor support and control components located thereon. <FIG> is a detailed view of the proprotor support and control components of <FIG>. As shown in <FIG>, in some examples, the boom <NUM> may comprise a tubular medial structure, and the distal arms <NUM>, 134r may be narrower tubular components coaxially mounted partially within the tubular medial structure. As shown in <FIG> and <FIG>, the aircraft <NUM> includes a gearbox <NUM>, which may be located within the proprotor tilt control housing <NUM> illustrated in <FIG>. As shown in <FIG>, the gearbox <NUM> includes master gears <NUM> independently driven by proprotor tilt servos <NUM>. The master gears <NUM> are positioned so as to mesh with slave gears <NUM>, which are coaxial with and rotationally fixed to inner ends of the distal arms <NUM>, 134r. Thus, the tilt servos <NUM> can tilt the proprotors <NUM>, 136r by rotating the master gears <NUM>.

As shown in <FIG> and <FIG>, a proprotor motor <NUM> is disposed at an outer end of each distal arm <NUM>, 134r of the boom <NUM>. Each proprotor motor <NUM> is rotationally fixed with respect to the corresponding distal arm <NUM>, 134r, such that the rotational axis of the proprotor motor <NUM> is perpendicular to the lateral axis along the distal arm <NUM>, 134r. Thus, rotation of a distal arm <NUM>, 134r about the lateral axis, under control of the proprotor tilt servos <NUM>, produces fore and aft tilting of the rotational axis of the proprotor motor <NUM>. Accordingly, each of the tilting operations described above with respect to the proprotors <NUM>, 136r may be achieved by actuating the proprotor tilt servos <NUM> individually or simultaneously.

<FIG> illustrate an example lower rotor hub <NUM> of an example aircraft such as the aircraft <NUM>. As described above, the lower rotor hub <NUM> serves as an attachment point for the upper rotor hub assembly <NUM> and the central rotor <NUM>. The lower rotor hub <NUM> is also configured to provide tilting and powered rotation of the central rotor <NUM>, as will now be described. The lower rotor hub <NUM> includes a rotor mount shaft <NUM> having one or more mounting pin holes <NUM> extending therethrough, a central rotor slave gear <NUM> meshed with a central rotor master gear <NUM>, a central rotor drive motor <NUM>, a tilt bearing <NUM>, and a central rotor tilt servo <NUM>.

The rotor mount shaft <NUM> serves as a mounting point for the central rotor <NUM>. The central rotor <NUM>, not shown in <FIG>, may be mounted to the aircraft <NUM> by coupling the upper rotor hub assembly <NUM> to the rotor mount shaft <NUM> and securing the upper rotor hub assembly <NUM> using one or more pins extending through mounting pin holes <NUM>. Pin-based mounting of the upper rotor hub assembly <NUM> allows the central rotor <NUM> to be easily and quickly attached to or detached from the aircraft <NUM> for transportation.

The central rotor slave gear <NUM> is coaxial with the rotor mount shaft <NUM> and is configured to transfer rotational motion of the central rotor master gear <NUM> to the rotor mount shaft <NUM> to drive the central rotor <NUM> during powered operation of the central rotor <NUM>. In some examples, the central rotor slave gear <NUM> is coupled to the rotor mount shaft <NUM> by a clutch mechanism and/or a one-way bearing (e.g., a one-way bearing <NUM>) such that rotational motion of the central rotor slave gear <NUM> in a first direction (e.g., clockwise) is transferred to the rotor mount shaft <NUM>, but the rotor mount shaft <NUM> is free to spin in the same direction (e.g., clockwise) when the central rotor slave gear <NUM> is not rotating or is rotating more slowly than the rotor mount shaft <NUM>. Thus, the central rotor drive motor <NUM> can power the central rotor <NUM> (e.g., during vertical flight and/or during the transition from vertical flight to forward flight) by turning the central rotor master gear <NUM>, which in turn causes the central rotor slave gear <NUM> and rotor mount shaft <NUM> to rotate.

The central rotor tilt servo <NUM> is configured to tilt the lower rotor hub <NUM> relative to the mast <NUM>. Actuation of the central rotor tilt servo <NUM> causes rotation of the lower rotor hub <NUM> about the tilt bearing <NUM>, which accommodates motion about a lateral axis perpendicular to the rotor mount shaft <NUM>.

<FIG> illustrates an example forward-most tilt position of the lower rotor hub <NUM>. In the tilt position illustrated in <FIG>, the rotor mount shaft <NUM> is substantially aligned with a vertical axis of the aircraft <NUM>, and a spinning central rotor <NUM> attached to the lower rotor mount <NUM> would produce a lifting force directly upward. The tilt position illustrated in <FIG> may be used, for example, during vertical flight while the proprotors <NUM>, 136r are in the VTOL/hover position illustrated in <FIG>.

<FIG> illustrates an example rear-most tilt position of the lower rotor hub <NUM>. In the tilt position illustrated in <FIG>, the rotor mount shaft <NUM> is tilted rearward approximately <NUM> degrees relative to vertical, such that a spinning central rotor <NUM> attached to the lower rotor mount <NUM> would rotate within a plane tilted approximately <NUM> degrees relative to a longitudinal axis of the aircraft <NUM>. The tilt position illustrated in <FIG> may be used, for example, during horizontal flight while the proprotors <NUM>, 136r are in the forward flight position illustrated in <FIG>. The central rotor tilt servo <NUM> tilts the lower rotor hub <NUM> between the positions illustrated in <FIG> and <FIG> when the aircraft <NUM> transitions from vertical flight to forward flight, or from forward flight to vertical flight, as described above.

<FIG> depict an example upper rotor hub assembly such as the upper rotor hub assembly <NUM> of the aircraft <NUM>. <FIG> depicts the upper rotor hub assembly <NUM> including two mounting members <NUM>, <NUM> coupleable to the lower rotor hub <NUM> of <FIG>, as well as an inner portion of a rotor blade <NUM> of the central rotor <NUM>. <FIG> and <FIG> depict the mounting members <NUM>, <NUM> separately, including an example rotor blade <NUM> mounted to the first mounting member <NUM>. <FIG> and <FIG> depict the upper rotor hub assembly <NUM> in a folded configuration.

As described above, the central rotor <NUM> of the aircraft <NUM> functions similarly to the rotor of a gyroplane when the aircraft <NUM> is in a forward flight mode. In contrast to helicopter rotors, in which blades may be hinged or otherwise able to move vertically and/or horizontally relative to the main rotor hub, gyroplane rotors perform most efficiently when the blades are rigidly fixed relative to the central hub and each pair of opposing blades remains symmetrically opposed. While not required, exact alignment of the blades may substantially improve performance. Thus, if compact transportation of the aircraft <NUM> is desired, it may be cumbersome to remove the blades <NUM> from the upper rotor hub assembly <NUM> for transport due to the time required to carefully align the blades <NUM> when reattaching them. As will now be described, the upper rotor hub assembly <NUM> is easily detachable from and attachable to the lower rotor hub <NUM>, and may be folded such that the entire central rotor <NUM> and upper rotor hub assembly <NUM> may be transported in a compact form while the blades <NUM> remain attached to the mounting members <NUM>, <NUM>.

The upper rotor hub assembly <NUM> includes a first mounting member <NUM> and a second mounting member <NUM>. The mounting members <NUM>, <NUM> each include mounting pin holes <NUM> configured to align with the mounting pin holes <NUM> of the lower rotor hub <NUM> of <FIG>. As shown in <FIG>, the mounting members <NUM>, <NUM> are shaped so as to nest together when both mounting members <NUM>, <NUM> are mounted to the lower rotor hub <NUM>. In addition, the mounting members <NUM>, <NUM> are shaped such that, when nested together in a mounted configuration, each mounting member <NUM>, <NUM> can teeter independently about the axis of its mounting pin. In some examples, the aircraft <NUM> can be operated with only two blades <NUM> by using only the first mounting member <NUM> or only the second mounting member <NUM>. Two-bladed operation is associated with less lift and less induced drag during forward flight, and may be desirable when the aircraft <NUM> is carrying a relatively light load and less lifting capacity is required; four-bladed operation is associated with more lift and more induced drag during forward flight, and may be desirable when the aircraft <NUM> is carrying a relatively heavy load and greater lifting capacity is required.

Each mounting member <NUM>, <NUM> includes two mounting brackets <NUM> disposed on opposite sides of the mounting member <NUM>, <NUM> about the center of the hub. Each mounting bracket <NUM> includes two or more mounting holes <NUM> spaced apart to align with corresponding mounting holes <NUM> within the rotor blades <NUM>. In some examples, more than two mounting holes <NUM>, <NUM>, such as three or more mounting holes, are provided in order to form a structurally robust connection between the rotor blades <NUM> and the mounting brackets <NUM>.

The mounting members <NUM>, <NUM> each include two hinges <NUM> disposed between the mounting brackets <NUM> and the central portion of the mounting members <NUM>, <NUM>. The hinges <NUM> allow the mounting brackets <NUM> and rotor blades <NUM> to be folded about the axis of the hinges <NUM> while the aircraft is not in flight. Each mounting bracket <NUM> is rigidly coupled to locking plates <NUM> having holes located to align with locking pin holes <NUM> of the central portion of the mounting members <NUM>, <NUM>. When a hinge <NUM> is in the fully extended position (e.g., for flight), the hinge <NUM> may be locked in the fully extended position by inserting a locking pin <NUM> through locking pin holes <NUM> and the adjacent locking plates <NUM>. The locking pins <NUM> may include a retaining mechanism such as spring-loaded retaining balls or the like, to prevent the locking pins <NUM> from pulling out during operation.

In some examples, instead of or in addition to being foldable, the mounting members <NUM>, <NUM> may also permit the rotor blades <NUM> to be removed from the mounting members <NUM>, <NUM> in a manner that retains the alignment of the rotor blades <NUM> when they are reinserted. For example, the rotor blades <NUM> may be slidably mounted along one or more rails disposed within the mounting brackets <NUM>. Each mounting bracket <NUM> may include a release button which, when depressed, permits a rotor blade <NUM> within the mounting bracket <NUM> to slide outward to be removed. In some examples, the mounting brackets <NUM> and/or the mounting members <NUM>, <NUM> may include a bayonet-type mounting system which maintains the appropriate alignment between opposing rotor blades <NUM>. Bayonet-type mounting systems will be described in greater detail with reference to <FIG> and <FIG>.

<FIG> and <FIG> illustrate the folded configuration of mounting members <NUM>, <NUM>. As shown in <FIG> and <FIG>, the removal of the locking pins <NUM> from the mounting members <NUM>, <NUM> allows the locking plates <NUM> to move away from the locking pin holes <NUM>, allowing the mounting members <NUM>, <NUM> to fold at the hinges <NUM> into a folded configuration. In the folded configuration, both mounting brackets <NUM> and rotor blades <NUM> attached to each mounting member <NUM>, <NUM> are substantially parallel. Thus, a two-bladed or four-bladed central rotor <NUM> can be transported within a relatively narrow rectangular container having a length only slightly longer than each of the rotor blades <NUM>. This folding configuration is substantially more efficient than transporting the central rotor <NUM> in a flight configuration, which would require a container having two perpendicular dimensions of at least twice the length of each rotor blade <NUM>. Moreover, in some cases it may be desirable to fold the central rotor <NUM> while the central rotor <NUM> remains attached to the aircraft <NUM>. The folding mechanism described herein permits the two or four rotor blades <NUM> to be folded upward such that the horizontal footprint of the aircraft <NUM> may be minimized while it is being stored on the ground.

<FIG> and <FIG> depict the upper rotor hub assembly <NUM> of <FIG> coupled to the lower rotor hub <NUM> of <FIG>. Each mounting member <NUM>, <NUM> is coupled to the lower rotor hub <NUM> by a mounting pin <NUM> passing through the mounting pin holes <NUM> of the mounting member <NUM>, <NUM> and the mounting pin holes <NUM> of the lower rotor hub <NUM>. Similar to the locking pins <NUM> that lock the mounting members <NUM>, <NUM> in the fully extended position, the mounting pins <NUM> may include a retaining mechanism such as spring-loaded retaining balls or the like, to prevent the mounting pins <NUM> from pulling out during operation.

<FIG> and <FIG> illustrate a further example configuration of an upper rotor hub assembly <NUM>. The example upper rotor hub assembly <NUM> of <FIG> and <FIG> further includes an upper rotor hub <NUM> which serves as a removable attachment point for the mounting members <NUM>, <NUM>. The mounting members <NUM>, <NUM> may be coupled to the upper rotor hub <NUM> such as by similar mounting pins <NUM>. In some examples, the mounting pins <NUM> coupling the mounting members <NUM>, <NUM> to the upper rotor hub <NUM> may be permanent or semi-permanent pins, rather than the easily removable pins of <FIG>.

As shown in <FIG>, the upper rotor hub <NUM> may in turn be connected to the rotor mount shaft <NUM> of the lower rotor hub by a further mounting pin <NUM>, which may similarly include spring-based or other retaining elements configured to be quickly releasable. Thus, the entire upper rotor mount hub assembly <NUM> may be folded and subsequently removed from the aircraft by removing a single mounting pin <NUM>.

<FIG> and <FIG> illustrate a further example configuration of an upper rotor hub assembly <NUM> in which a bayonet-type mounting system allows the rotor blades <NUM> to be attached and detached from the upper rotor hub assembly <NUM>. Similar to the configuration of <FIG> and <FIG>, the upper rotor hub assembly <NUM> of <FIG> and <FIG> includes two mounting members <NUM>, <NUM> attached to an upper rotor hub <NUM>. Each rotor blade <NUM> is mounted to a mounting bracket <NUM> by two, three, or more fasteners. Each mounting bracket <NUM> is fixed to a mounting body <NUM>, a portion or all of which may be integrally formed with the mounting bracket <NUM>. A blade attachment pin hole <NUM> extends laterally through the mounting body <NUM>.

Each opposing end of each mounting member <NUM>, <NUM> includes a mounting body opening <NUM> sized and shaped to receive a mounting body <NUM>. Additional blade attachment pin holes <NUM> extend laterally through the sides of the mounting members <NUM>, <NUM>. Thus, as illustrated by the transition from <FIG>, each rotor blade <NUM> may be mounted by sliding the mounting body <NUM> into a mounting body opening <NUM> until the blade attachment pin holes <NUM> are substantially aligned with blade attachment pin holes <NUM>. A blade attachment pin <NUM> may then be inserted through the blade attachment pin holes <NUM>, <NUM> to secure the rotor blade <NUM> to the upper rotor hub assembly <NUM>. Dismounting of the blades, such as for storage, transport, etc., may be accomplished by removing the blade attachment pin <NUM> from the blade attachment pin holes <NUM>, <NUM> and subsequently sliding the mounting body <NUM> out of the mounting body opening <NUM>.

As described above, it is typically desirable for the rotor to have a positive blade pitch when operating in vertical powered flight, such as in a hover or VTOL phases of flight. In contrast, when autorotation is used, such as in forward flight of a gyroplane, it is desirable for the rotor blades to have a flat or zero pitch, or a substantially less positive pitch than in vertical flight. Accordingly, some embodiments of the present technology include rotor assemblies configured to selectively change the pitch angle of the rotor blades while maintaining the teetering motion desirable for low-pitch forward flight. Advantageously, the embodiments disclosed herein accomplish blade pitch control without requiring the weight and complexity of a swashplate as is typically utilized for blade pitch control in helicopters.

<FIG> depicts an example embodiment of a rotor hub assembly including a teetering pivot point and synchronization linkages for controlling the pitch of the rotor blades. <FIG> and <FIG> illustrate side and cross-sectional views of the rotor hub assembly of <FIG> in a zero pitch configuration. <FIG> and <FIG> illustrate side and cross-sectional views of the rotor hub assembly of <FIG> in a positive pitch configuration. <FIG> depicts an exploded view of the rotor hub assembly of <FIG>. The rotor assembly illustrated in <FIG> may be implemented in conjunction with any of the aerial vehicles disclosed herein.

In order for a gyrocopter rotor assembly to generate upward thrust, the pitch of the blades is increased from zero or approximately zero to a selected pitch angle greater than zero. Additionally, the rotor is turned at an RPM that will generate enough thrust to lift the weight of the aircraft and payload. <FIG> and <FIG> illustrate the assembly with the rotor blades being at a zero-degree pitch, and <FIG> and <FIG> illustrate the assembly with the rotor blades being at a higher pitch that creates lift.

As shown in <FIG>, <FIG>, and <FIG>, the rotor head assembly includes a teetering pivot point attached to the rotating shaft, such that the rotor head can teeter freely during forward flight. The rotor head assembly further includes synchronization linkages connecting the two opposing arms of the rotor head that holds the two rotor blades. These linkages synchronize the movement of the two arms as they slide out and twist to create the desired pitch.

Each arm of the rotor assembly includes an outer cylinder configured to retain a removable rotor blade, as described in greater detail with reference to <FIG> and <FIG>, and an inner cylinder sized and shaped to at least partially fit within the outer cylinder. A biasing element such as a gas spring is disposed between an end backstop of the inner cylinder and an inward-facing surface of the outer cylinder arm. The outer cylinder arm is connected on both ends to the synchronization linkages via arms <NUM> and <NUM> extending through arm apertures in the outer surface of the outer cylinder. Each arm passes through the two arm apertures of the corresponding outer cylinder and the two rotor orientation tracks of the corresponding inner cylinder, and is coupled at each end to an outer end of a synchronization linkage.

Advantageously, the rotor assembly mechanism of <FIG> can provide for automatic adjustment of blade pitch based on the rotational speed of the rotor assembly, without requiring a servo or other control mechanism to adjust the blade pitch. For example, the gas spring and associated components can cause the blade pitch to increase automatically at higher RPM, and to decrease automatically at low RPM. The gas spring is rated to collapse at a certain pound force. Thus, as the rotational speed of the rotor increases in a powered vertical flight mode, the corresponding increase in centrifugal force causes the gas spring to at least partially collapse, letting the outer cylinder slide outwards. As the outer cylinder slides outward, the slope of the rotor orientation tracks causes the outer cylinder and attached rotor blade to twist, creating a desired pitch for the blade. The synchronization linkages simultaneously slide outwards and twist to the desired pitch for lift generation, and maintain an equal or substantially equal blade pitch between the two blades. As the RPM is subsequently reduced, the outer cylinders slide inwards through the assistance of the two gas springs, bringing the blade angle back to zero as arms <NUM> and <NUM> slide within the rotor orientation tracks, allowing the gyrocopter to fly forward with the aid of the pusher motor/propeller assembly.

In some embodiments, the gas spring may be configured to collapse when a predetermined outward force (e.g., radially outward from the center of the rotor assembly toward the blade) is applied. The predetermined force may be selected, based at least partially on the mass of the outer cylinder and blade, such that the outer cylinder collapses and increases the rotor blade pitch at a predetermined range of rotational speeds. In one particular example, the predetermined force is selected such that a lower RPM range, such as <NUM>-<NUM> RPM, does not create sufficient centrifugal force to collapse the gas spring outward, while a higher RPM range, such as above <NUM> RPM, creates sufficient centrifugal force to cause the gas spring to remain collapsed. The aircraft may thus be controlled to operate with the main rotor turning at <NUM> RPM or slower while in horizontal flight, and with the main rotor turning at <NUM> RPM or faster while in vertical or hovering flight. In some embodiments, the aircraft may be configured to avoid operating for extended periods with the main rotor turning at speeds in an intermediate or safety RPM range (e.g., between <NUM> RPM and <NUM> RPM in the particular example above) at which the centrifugal force created by the rotor blades and outer cylinders may be great enough to partially collapse the gas spring, but may not be sufficient to fully collapse the gas spring.

Consistent with the automatic control of blade pitch based on rotor RPM, it may be desirable to increase and decrease the rotational speed of the rotor on command. In addition, it may be desirable to increase or decrease rotor RPM during flight in order to achieve desirable or optimized flight characteristics. <FIG> illustrates an example gearing system for increasing the RPM of the upper rotor hub assembly of <FIG>. <FIG> is a cross-sectional view illustrating a braking system for reducing the rotational speed of the upper rotor hub assembly of <FIG>.

As shown in <FIG>, rotational speed of the rotor may be increased by a motor via a large gear having a one-way bearing, as described elsewhere herein. When that motor is powered and turns the rotor head and blades attached to the rotor head, it generates a significant torque that if not countered, will spin the aircraft uncontrollably. In order to avoid undesirable spinning, the two proprotors that are located on either side of the vehicle (e.g., as shown in <FIG>) at a specific CG (center of gravity) location, can operate independently at different speeds so as to offset the torque created by the rotor head assembly. The rotational speeds of the proprotors are variable in order to match the counter torque force required based on the torque created by the rotor head assembly. In some embodiments, depending on the amount of torque generated by the rotor head assembly, one proprotor (e.g., the left proprotor) may operate at up to full speed while the other proprotor operates at a lower RPM or even at a full stop in order to fully counter the torque of the rotor head assembly. In some embodiments, the other proprotor (e.g., the right proprotor) may even be rotated <NUM> degrees and operated at a suitable RPM to further provide counter torque. When the aircraft transitions to forward flight, the proprotors can transition to turn at the same RPM so as to provide stability and maneuverability during the horizontal phase of flight.

Referring now to <FIG>, in order to slow down or control the rotor RPM, a disk brake is located at the bottom of the rotor shaft. The disk brake is coupled with a bracket that is part of the pivoting assembly bracket and which holds the brake pads, calipers, and brake servo. The calipers that hold the brake pads are activated by the brake servo and are activated at any time in order to slow down the rotor or assist in stopping it completely once the aircraft has landed safely. This bracket assembly allows the rotor head to move forwards and backwards as required for the desired flight mode.

<FIG> illustrate an example aircraft including a tiltable empennage. In some examples, the empennage, including the horizontal stabilizer and one or two rudders, are turned <NUM> degrees downwards during vertical flight, thereby streamlining the airflow created by the downwards thrust from the rotating blades of the main rotor. The empennage may be rotatable about a lengthwise axis of the horizontal stabilizer (e.g., a lateral axis with regard to the aircraft). <FIG> and <FIG> illustrate the empennage in an upright position, such as for forward flight. <FIG> illustrates the empennage in a lowered position, tilted approximately <NUM> degrees forward or backward such that the horizontal stabilizer is oriented substantially within a vertical plane. Once the tail assembly is in the lowered position, it has two additional functions: First, by changing the angle of the stabilizer from <NUM> degrees plus/minus, it will move the gyrocopter forwards or backwards in a controlled manner. Second, the rudders allow for additional counter torque fine tuning and being able to turn the aircraft left or right. The down turned wings have the same capability but have the primary counter torque function. Moreover, the vertical orientation of the horizontal stabilizer in the lowered position reduces interference with the downward airflow created by the rotor, improving vertical and hover flight efficiency. The rotation of the tiltable empennage between upright and lowered positions may be actuated by an empennage tilt servo disposed within the aircraft.

<FIG> illustrate a further example of an aerial vehicle including a single proprotor and pivotable wings configured to operate as control surfaces. The aerial vehicle of <FIG> includes a central rotor and a single proprotor mounted to the mast so as to provide centerline thrust for the aerial vehicle. Independently pivotable wings are mounted at the sides of the mast. It will be understood that the single proprotor configuration of <FIG> may be implemented with any of the aerial vehicle examples disclosed herein.

As shown in <FIG> and <FIG>, in horizontal forward flight, the two wings have substantially the same horizontal orientation such that the wings are streamlined for forward flight. In some examples, the wings may further be used for flight control functionality in forward flight, such as to provide a rolling or pitching moment.

As shown in <FIG> and <FIG>, the two wings can be differentially positioned so as to provide a counter torque moment while the main rotor is powered in vertical or hovering flight phases. In this counter torque position, the left wing is pivoted by more than <NUM> degrees and the right wing is pivoted by less than <NUM> degrees, such that the left wing deflects the rotor downwash forward and the right wing deflects the rotor downwash aft. Thus, the counter torque configuration of <FIG> and <FIG> creates a counterclockwise torque that counters the clockwise torque created by the powered rotor.

While certain embodiments have been described, these embodiments have been presented by way of example only, and are not intended to limit the scope of the disclosure. Indeed, the novel methods and systems described herein may be embodied in a variety of other forms. The accompanying claims are intended to cover such forms or modifications as would fall within the scope of the invention. Accordingly, the scope of the present invention is defined only by reference to the appended claims.

Features, materials, characteristics, or groups described in conjunction with a particular aspect, embodiment, or example are to be understood to be applicable to any other aspect, embodiment or example described in this section or elsewhere in this specification unless incompatible therewith.

Furthermore, certain features that are described in this disclosure in the context of separate implementations can also be implemented in combination in a single implementation.

Claim 1:
A rotor assembly for an aircraft comprising:
a plurality of rotor blades (<NUM>); and
an upper rotor hub assembly (<NUM>) including at least one mounting member (<NUM>, <NUM>) each mounting member comprising:
a central section comprising a pivot coupler configured to allow the mounting member and rotor blades to teeter relative to the aircraft;
first and second inner cylinders extending from the central section, the first and second inner cylinders comprising apertures defining a track;
two mounting brackets (<NUM>) disposed at opposite ends of the central section, each mounting bracket fixedly coupled to one of the plurality of rotor blades, the mounting brackets comprising first and second outer cylinders that slidingly receive the first and second inner cylinders, the outer cylinders comprising pin-receiving apertures; and
a blade pitch adjustment linkage coupled to the mounting brackets and the central section, the blade pitch adjustment linkage configured to automatically adjust a blade pitch of the rotor blades based on a rotational velocity of the rotor blades and to synchronize the blade pitch of the rotor blades,
wherein the blade pitch adjustment linkage comprises pins extending through the apertures in the outer and inner cylinders such that when the outer cylinders slide relative to the inner cylinders, the pins slide through the tracks, thereby causing rotation of the outer cylinders relative to the inner cylinders.