Patent Publication Number: US-11661182-B2

Title: Aerial vehicle

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This application is a continuation of U.S. application Ser. No. 17/461,719, filed Aug. 30, 2021, entitled “AERIAL VEHICLE,” which is a continuation-in-part of International PCT Application No. PCT/US2020/046240, filed Aug. 13, 2020, entitled “AERIAL VEHICLE,” which claims the benefit of U.S. Provisional Application Ser. No. 62/886,578, filed Aug. 14, 2019, entitled “AERIAL VEHICLE,” U.S. Provisional Application Ser. No. 62/896,257, filed Sep. 5, 2019, entitled “AERIAL VEHICLE,” and U.S. Provisional Application Ser. No. 63/018,848, filed May 1, 2020, entitled “AERIAL VEHICLE,” all of which are incorporated by reference in their entirety and for all purposes. 
    
    
     FIELD 
     The present disclosure relates generally to manned or unmanned aircraft, and more particularly to aircraft capable of vertical and horizontal flight. 
     BACKGROUND 
     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 100% 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&#39;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 mounted 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. 
     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. 
     SUMMARY 
     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. Without limiting the scope of this disclosure, its more prominent features will now be discussed briefly. 
     In a first aspect, an aircraft comprises a fuselage, a mast extending upward from the fuselage and fixed to the fuselage at a lower end of the mast, a rotor rotatably coupled to an upper end of the mast opposite the lower end, a rotor motor disposed at the upper end of the mast and configured to cause rotation of the rotor in a first direction, a lateral boom coupled to an intermediate section of the mast between the upper end and the lower end, a first powered proprotor disposed at a first end of the lateral boom, a second powered proprotor disposed at a second end of the lateral boom opposite the first end, a first proprotor tilt servo configured to control an orientation the first powered proprotor between at least a horizontal tilt angle and a vertical tilt angle, and a second proprotor tilt servo configured to control an orientation of the second powered proprotor between at least the horizontal tilt angle and the vertical tilt angle independent of the orientation of the first powered proprotor. 
     In some embodiments, the first powered proprotor is powered by a first proprotor motor, the second powered proprotor is powered by a second proprotor motor, and rotational speeds of the first proprotor motor and the second proprotor motor are independently variable. In some embodiments, the first and second proprotor motors are configured to counteract torque effects of the rotor motor during vertical or hovering flight by powering the first and second proprotors at different speeds. In some embodiments, an output of the rotor motor is rotationally coupled to the rotor by a one-way bearing that transmits torque from the rotor motor to the rotor when the rotor motor is activated and allows the rotor to turn freely in the first direction when the rotor motor is deactivated. In some embodiments, the first and second proprotor tilt servos are configured to control the orientations of the first and second powered proprotors within a range of tilt angles of greater than 90 degrees. In some embodiments, the range of tilt angles is approximately 150 degrees. In some embodiments, the first and second powered proprotors are tiltable between a first extreme tilt angle of at least 25 degrees aft of vertical and a second extreme tilt angle of at least 25 degrees below horizontal. In some embodiments, the first extreme tilt angle is at least 30 degrees aft of vertical, and the second extreme tilt angle is at least 30 degrees below horizontal. In some embodiments, the aircraft further comprises a rotor tilt servo configured to tilt the rotor between at least a first tilt angle in which the rotor spins about a vertical axis of the aircraft and a second tilt angle in which the rotor spins about an axis angled at approximately 20 degrees relative to the vertical axis of the aircraft. In some embodiments, the first and second proprotor tilt servos are configured to counteract torque effects of the rotor motor during vertical or hovering flight by differentially tilting the first and second powered proprotors relative to a vertical axis of the aircraft. In some embodiments, the aircraft is configured to fly in a plurality of flight configurations including a vertical flight configuration in which the first and second powered proprotors are disposed within 30 degrees of vertical and the rotor is driven by the rotor motor, and a horizontal flight configuration in which the first and second powered proprotors are disposed within 30 degrees of horizontal and the rotor turns by free autorotation. In some embodiments, the rotor comprises rotor blades configured to automatically adjust to a positive blade pitch in the vertical flight configuration and to a flat blade pitch in the horizontal flight configuration. In some embodiments, the aircraft further comprises a tiltable empennage including a horizontal stabilizer, the tiltable empennage configured to rotate about a lengthwise axis of the horizontal stabilizer. In some embodiments, the tiltable empennage is configured to rotate to a lowered position during vertical flight such that the horizontal stabilizer is aligned with a downwash created by the rotor when the rotor motor is engaged. 
     In a second aspect, an aircraft comprises a rotor comprising a plurality of rotor blades, a lower rotor hub, and an upper rotor hub assembly including at least one mounting member. The lower rotor hub comprises a rotor mount shaft extending along an axis of rotation of the rotor, and at least one mounting pin hole extending through the rotor mount shaft. Each mounting member comprises a central section having two mounting pin holes extending therethrough such that the central section can be coupled to the lower rotor hub by inserting a mounting pin through the two mounting pin holes of the central section, and two mounting brackets disposed at opposite ends of the central section, each mounting bracket fixedly coupled to one of the plurality of rotor blades. 
     In some embodiments, the upper rotor hub assembly further comprises a blade pitch adjustment linkage configured to automatically adjust a blade pitch of the rotor blades based on a rotational velocity of the rotor. In some embodiments, the blade pitch adjustment linkage comprises at least one biasing element configured to cause an increase in the blade pitch when subjected to an increased centrifugal force associated with an increase in the rotational velocity of the rotor. In some embodiments, the biasing element comprises a gas spring coupled to the central section and one of the two mounting brackets. In some embodiments, each mounting bracket of each mounting member is coupled to the central section by a hinge allowing the mounting member to be folded between an extended configuration and a folded configuration, each mounting bracket further comprising a plurality of locking pin holes configured to lock the mounting member in the extended configuration when a locking pin is disposed within the locking pin holes. In some embodiments, each mounting member can teeter about the mounting pin. In some embodiments, the aircraft comprises at least four blades and at least two mounting members, each mounting member being coupled to two oppositely disposed rotor blades of the plurality of rotor blades, and each mounting member is shaped to permit at least one other mounting member to teeter independently. In some embodiments, each rotor blade is slidably attachable and detachable from the mounting brackets. In some embodiments, each mounting bracket of each mounting member is attached to a mounting body sized and shaped to fit within a mounting body opening of the mounting member, and the mounting body is securable to the mounting member by inserting a blade attachment pin through pin holes in the mounting member and the mounting body. In some embodiments, the upper rotor hub assembly further comprises an upper rotor mount hub comprising a tubular shaft at least partially surrounding the rotor mount shaft of the lower rotor mount hub, wherein each mounting member is coupled to the upper rotor mount hub by at least one mounting pin, and wherein the tubular shaft is detachably coupled to the lower rotor mount shaft by at least one mounting pin. 
     In a third aspect, a rotor assembly for an aircraft comprises a plurality of rotor blades and an upper rotor hub assembly including at least one mounting member. Each mounting member comprises a central section; two mounting brackets disposed at opposite ends of the central section, each mounting bracket fixedly coupled to one of the plurality of rotor blades; 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. 
     In some embodiments, the blade pitch adjustment linkage comprises at least one biasing element configured to cause an increase in the blade pitch when subjected to an increased centrifugal force associated with an increase in the rotational velocity of the rotor blades. In some embodiments, the biasing element comprises a gas spring coupled to the central section and one of the two mounting brackets. In some embodiments, the blade pitch adjustment linkage includes at least one synchronization linkage coupled to the mounting brackets and configured to synchronize the blade pitch of the rotor blades. 
     In a fourth aspect, a rotor assembly for an aircraft comprises a first and second rotor blade, and a mounting member to which the first and second rotor blades are coupled in an opposed arrangement. The mounting member comprises a central section comprising a pivot coupler configured to allow the mounting member and rotor blades to teeter relative to the aircraft; two mounting brackets disposed at opposite ends of the central section, each mounting bracket fixedly coupled to first and second rotor blades; and a blade pitch linkage coupled to the mounting brackets and the central section, the blade pitch linkage configured to synchronize a pitch angle of the first and second rotor blades. 
     In some embodiments, the rotor assembly comprises first and second inner cylinders extending from the central section, wherein the mounting brackets comprise first and second outer cylinders that slidingly receive the first and second inner cylinders, wherein the outer cylinders comprise pin-receiving apertures and the inner cylinders comprise apertures defining a track, wherein the blade pitch linkage comprises pins extending through the apertures in the first and second 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. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       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. 
         FIG.  1    depicts an isometric view of an aircraft in accordance with an example embodiment of the present technology. 
         FIG.  2    depicts a side view of the aircraft of  FIG.  1   . 
         FIG.  3    depicts a rear isometric view of the aircraft of  FIGS.  1  and  2   . 
         FIG.  4    depicts a top front view of the aircraft of  FIGS.  1 - 3   . 
         FIG.  5    schematically illustrates an example tilting range of proprotors with respect to the lateral boom of an example aircraft. 
         FIG.  6    schematically illustrates example VTOL/hover and forward flight tilt positions of the proprotors. 
         FIG.  7    depicts a medial portion of the lateral boom of an example aircraft including proprotor support and control components located thereon. 
         FIG.  8    is a detailed view of certain proprotor support and control components of  FIG.  7   . 
         FIG.  9    depicts medial and lateral portions of the lateral boom of the example aircraft of  FIGS.  7  and  8    including proprotor support and control components located thereon. 
         FIG.  10    depicts a lateral portion of the lateral boom of the example aircraft of  FIGS.  7 - 9    including proprotor support and control components located thereon. 
         FIG.  11    depicts a side view of a lower rotor hub of an example aircraft. 
         FIG.  12    depicts a rear view of the lower rotor hub of the example aircraft of  FIG.  11   . 
         FIG.  13    depicts a top perspective view of the lower rotor hub of the example aircraft of  FIGS.  11  and  12   . 
         FIG.  14    depicts a side view of the lower rotor hub of the example aircraft of  FIGS.  11 - 13    in a vertical configuration. 
         FIG.  15    depicts a side view of the lower rotor hub of the example aircraft of  FIGS.  11 - 14    in a tilted configuration. 
         FIG.  16    depicts an upper rotor hub assembly including two mounting members coupleable to the lower rotor hub of  FIGS.  11 - 15   . 
         FIGS.  17  and  18    depict a rotor blade mounted to a blade arm of a mounting member in accordance with an example embodiment. 
         FIGS.  19  and  20    depict the upper rotor hub assembly in a folded configuration. 
         FIG.  21    depicts a partial isometric view of the upper rotor hub assembly of  FIGS.  16 - 20    coupled to the lower rotor hub of  FIGS.  11 - 15   . 
         FIG.  22    depicts a partial top view of the upper rotor hub assembly of  FIGS.  16 - 20    coupled to the lower rotor hub of  FIGS.  11 - 15   . 
         FIG.  23    depicts a perspective view of a portion of an upper rotor hub assembly in accordance with some embodiments. 
         FIG.  24    depicts a side view of the portion of an upper rotor hub assembly of  FIG.  23    coupled to the lower rotor hub of an example aircraft. 
         FIGS.  25  and  26    depict an example embodiment of an upper rotor hub assembly with a bayonet-type blade mounting system. 
         FIG.  27    depicts an example embodiment of a rotor hub assembly including a teetering pivot point. 
         FIGS.  28  and  29    illustrate side and cross-sectional views of the rotor hub assembly of  FIG.  27    in a zero pitch configuration. 
         FIGS.  30  and  31    illustrate side and cross-sectional views of the rotor hub assembly of  FIGS.  27 - 29    in a positive pitch configuration. 
         FIG.  32    depicts an exploded view of the rotor hub assembly of  FIGS.  27 - 31   . 
         FIG.  33    illustrates an example gearing system for turning the upper rotor hub assembly of  FIGS.  27 - 32   . 
         FIG.  34    is a cross-sectional view illustrating a braking system for reducing the rotational speed of the upper rotor hub assembly of  FIGS.  27 - 32   . 
         FIGS.  35 - 37    illustrate an example aircraft including a tiltable empennage. 
         FIGS.  38 - 41    illustrate an example aircraft including a single proprotor and pivotable wings configured to operate as control surfaces. 
         FIGS.  42  and  43    illustrate a further example configuration of an aircraft in accordance with the present technology. 
         FIGS.  44  and  45    illustrate a further example configuration of a rotor head assembly in accordance with the present technology. 
     
    
    
     DETAILED DESCRIPTION 
     Generally described, embodiments of the present disclosure provide 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. 
       FIGS.  1 - 4    illustrate an example aircraft  100  configured for in-flight transition between vertical and horizontal flight modes. The aircraft  100  includes a fuselage  102 , a mast  120  extending generally upward from the fuselage  102 , and an empennage  150  disposed aft of the fuselage  102 . A tiltable central rotor  140  is rotatably coupled at an upper end of the mast  120 . Tiltably mounted proprotors  136   l,    136   r  are rotatably coupled at opposing ends of a boom  132  extending laterally from the mast  120 . 
     The fuselage  102  is a body section of the aircraft  100  and may include an interior volume sized and shaped to hold a payload. For example, the interior volume of the fuselage  102  may be used to contain one or more items being transported by the aircraft  100 . In some embodiments, the fuselage  102  may contain one or more reconnaissance or 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  102  may further contain at least some of the control systems for the aircraft  100 , such as motor, control surface, or tilt servo controllers, autopilot systems, and the like. For example, in some embodiments, an autopilot system may be connected to proprotor speed and tilt control circuitry such that the autopilot can directly control the speed, direction of rotation, and/or tilt of the proprotors  136   l,    136   r  as pitch, roll, and/or yaw control devices. If the aircraft  100  is configured for operation as a remotely piloted unmanned aerial vehicle (UAV) or drone, the fuselage  102  may also include a communication system to receive control commands from a remote pilot. An undercarriage  104  may be disposed on a side or bottom portion of the fuselage  102  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  102  and the undercarriage  104  may comprise any suitably rigid or semi-rigid materials such as metal, plastic, carbon fiber, wood, fiberglass, etc. 
     The mast  120  extends generally upward from the fuselage  102  and supports the central rotor  140  disposed at an upper end of the mast  120 . In some embodiments, such as in the embodiment illustrated in  FIGS.  1 - 4   , the mast also includes a forward tilt such that the mast extends upward at a forward angle from the fuselage  102 . As shown in  FIG.  2   , the forward tilt of the mast  120  may advantageously allow the rotational axis of the central rotor  140  to align with the center of gravity of the fuselage  102 . In addition, when the proprotors  136   l,    136   r  are tilted upward for vertical/hover flight, the central rotor  140 , the longitudinal location of the proprotors  136   l,    136   r,  and the center of gravity of the fuselage  102  are all aligned with a common center of gravity for improved stability in hover or vertical flight. The mast  120  may also serve as a central attachment point for other components of the aircraft  100 , including the boom  132 , empennage  150 , and lower rotor hub  300 . The mast  120  may further include an interior volume in which one or more aircraft components may be disposed. In some embodiments, energy storage media such as batteries, hydrogen storage, or the like, may be contained within the mast  120 . Placement of batteries or hydrogen storage within the mast  120  may advantageously keep such relatively heavy components close to the center of gravity of the aircraft  100 , resulting in improved stability. 
     The central rotor  140  is rotatably mounted to a top portion of the shaft  120  via a lower rotor hub  300 . Rotor blades  142  are fixed to an upper rotor hub assembly  400  coupled to the lower rotor hub  300 . The rotor blades  142  may have an airfoil profile configured to enhance the production of lift while the central rotor  140  is spinning. Although the rotor  140  of the example aircraft of  FIGS.  1 - 4    has four rotor blades  142 , it will be understood that other numbers of rotor blades  142  may be included. For example, in some embodiments, the rotor  140  may have 2, 3, 4, 5, 6, or more rotor blades  142 . 
     The lower rotor hub  300  is tiltably mounted at the top of the mast  120 , for example, on a central rotor control housing  122 . The central rotor control housing  122  may include one or more servos configured to provide at least fore and aft tilting of the central rotor  140  (e.g., by tilting the lower rotor hub  300 ). In some embodiments, the central rotor control housing  122  further includes one or more servos configured to provide lateral tilting of the central rotor  140 . The central rotor control housing  122  can also include a rotor motor configured to turn the central rotor  140  during some phases of flight (e.g., during vertical and/or transitional flight modes), as described in greater detail below. In some embodiments, the rotor motor may be configured to turn the central rotor  140  at rotational speeds up to typical gyroplane rotor rotational speeds (e.g., up to approximately 200-600 rpm), which may be slower than typical helicopter rotor rotational speeds (e.g., approximately 400-1500 rpm or more). The rotor motor may be rotationally coupled to the central rotor  140  by a clutch and/or a one-way bearing such that, during forward flight, the central rotor  140  can rotate faster than the rotor motor, and such that the central rotor  140  can continue rotating by autorotation when the rotor motor is not turning. 
     The proprotors  136   l,    136   r  are configured to provide lift and/or forward thrust, depending on the tilt of the proprotors  136   l,    136   r.  The proprotors  136   l,    136   r  are tiltably mounted at opposite ends of the boom  132 . In some embodiments, the boom  132  includes distal arms  134   l  and  134   r,  which are individually pivotable along the lateral axis of the boom, such that each proprotor  136   l,    136   r  is independently tiltable by pivoting the distal arms  134   l,    134   r  of the boom  132 . Servos and/or other actuators disposed within a proprotor tilt control housing  130  may control pivoting of the distal arms  134   l,    134   r.  As described in greater detail below, each proprotor  136   l,    136   r  can be independently powered by left and right proprotor motors and may be operable at different relative speeds for enhanced maneuverability. For example, as described above, the speed, direction of rotation, and/or tilt of each proprotor  136   l,    136   r  may be independently controlled such as by an autopilot or other control system such that the proprotors  136   l,    136   r  can function as control devices to control roll, pitch, and/or yaw of the aircraft  100  without requiring conventional control surfaces. In another example, attitude of the aircraft may be controlled by controlling the thrust of the proprotors  136   l,    136   r  together in coordination with tilt of the proprotors  136   l,    136   r.  The speeds of both proprotors  136   l,    136   r  may further be adjusted collectively so as to propel the aircraft forward at a range of desired speeds. In some embodiments, the proprotors  136   l,    136   r  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  150  is mounted at a rear portion of the aircraft  100  and includes a horizontal stabilizer  152  and vertical stabilizers  154 . A longitudinal tail boom  156  may fix the empennage  150  to the mast  120  or fuselage  102 . The horizontal stabilizer  152  and vertical stabilizers  154  provide stability to the aircraft, primarily during horizontal flight. In some embodiments, the empennage  150  may include one or more control surfaces such as an elevator disposed on the horizontal stabilizer  152  and/or rudders disposed on the vertical stabilizers  154 . In embodiments having rudders on the vertical stabilizers  154 , the vertical stabilizers  154  may be placed at a location within the slipstreams of the proprotors  136   l,    136   r  to increase rudder effectiveness at low airspeeds. However, it will be understood that in various embodiments, the need for control surfaces may be eliminated by the use of independently controllable and tiltable proprotors  136   l,    136   r  as described above. 
     With reference to  FIGS.  5  and  6   , and with continued reference to  FIGS.  1 - 4   , example flight control systems and methods will now be described.  FIG.  5    schematically illustrates an example tilting range of proprotors  136   l,    136   r  with respect to the lateral boom  132  of the aircraft  100 .  FIG.  6    schematically illustrates example VTOL/hover and forward flight tilt positions of the proprotors. Each of the proprotors  136   l,    136   r  may have a range of tilt angles of at least 90 degrees, and in some cases up to  150  degrees or more. For example, proprotors  136   l,    136   r  may be tiltable to a VTOL/hover position or range of positions, in which each proprotor  136   l,    136   r  is substantially aligned about a vertical or z-axis to produce an upward lifting force. Proprotors  136   l,    136   r  may further be tiltable to a forward flight or horizontal flight position or range of positions, in which each proprotor  136   l,    136   r  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  136   l,    136   r  may be independently tiltable within a range of approximately 15 degrees, 20 degrees, 25 degrees, 30 degrees, or more, such as for maneuvering and/or stability as described below. In some embodiments, the proprotors  136   l,    136   r  can be tilted to any tilt angle between 30 degrees below horizontal to 30 degrees aft of vertical, for a full tilt angle range of approximately 150 degrees. 
     Vertical flight modes, such as VTOL, hovering, may be achieved with the proprotors  136   l,    136   r  in a VTOL/hover position, such as wherein the axis of rotation of each of the proprotors  136   l,    136   r  is within approximately 30 degrees of vertical. In the VTOL/hover position, the spinning proprotors  136   l,    136   r  primarily generate an upward lifting force. In addition, the central rotor  140  may also be used during vertical flight modes. The central rotor  140  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  100  and produce an additional lifting force in addition to the lift provided by the proprotors  136   l ,  136   r.  For example, turning the central rotor  140  at a relatively low rotational speed (e.g., a typical gyroplane rotor rotational speed) can be sufficient to significantly stabilize the aircraft  100 . However, in addition to stabilizing the aircraft  100 , the central rotor  140  when powered may also create torque effects and gyroscopic precession. In some embodiments, the proprotors  136   l,    136   r  may be differentially controlled, and/or may be configured to rotate in a direction opposite the rotation of the central rotor  140 , to counteract the torque effect of the central rotor  140 . For example, the torque generated by a clockwise turning central rotor  140  may tend to cause the body of the aircraft  100  to spin counterclockwise. To counteract this torque effect, the left proprotor  136   l  may be tilted slightly forward of vertical while the right proprotor  136   r  is tilted slightly aft of vertical, such that the proprotors  136   l,    136   r  exert a clockwise torque on the body of the aircraft  100  (e.g., a yawing moment) that counteracts the torque effect of the turning central rotor  140 . In some embodiments, a yawing moment may be created by varying the speeds and/or rotational directions of the proprotors  136   l,    136   r . For example, rotating proprotor  136   l  faster than proprotor  136   r  creates a yawing moment to the right, and rotating proprotor  136   r  faster than proprotor  136   l  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  136   l,    136   r  may further be used to maneuver the aircraft  100 . As described above, differential tilting, motor speed, or rotational direction of the proprotors  136   l,    136   r  may be used to produce a yawing moment for maneuverability about the vertical axis. Differential rotation speed, rotational direction, and/or tilt of the proprotors  136   l,    136   r  may be used to produce a rolling moment for maneuverability about the longitudinal axis. For example, powering the left proprotor  136   l  at a higher rotational speed relative to the right proprotor  136   r  produces a right or clockwise rolling moment; powering the right proprotor  136   r  at a higher rotational speed relative to the left proprotor  136   l  produces a left or counterclockwise rolling moment. Simultaneous tilting of the proprotors  136   l,    136   r  forward or aft of vertical, and/or tilting the central rotor  140  forward or aft, produces a pitching moment for maneuverability about the lateral axis. Rolling or pitching the aircraft  100  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  136   l,    136   r  in a forward flight position, such as wherein the axis of rotation of each proprotor  136   l,    136   r  is within approximately 30 degrees of horizontal. In the forward flight position, the spinning proprotors  136   l,    136   r  primarily generate forward thrust substantially parallel to a direction of flight. During forward flight, the central rotor  140  may be unpowered and turns in free autorotation. Preferably, the central rotor  140  has an upward tilt (e.g. between approximately 1 degree and 20 degrees or more with the higher side of the central rotor  140  oriented toward the direction of flight). In some embodiments, the central rotor  140  is tilted at an angle of approximately 1 to 20 degrees in forward flight. Thus, in forward flight, the aircraft  100  performs substantially as a gyroplane, with the powered proprotors  136   l,    136   r  generating thrust and the autorotating central rotor  140  providing lift. The empennage  150  provides directional stability during forward flight. 
     In various embodiments, the empennage  150  may or may not include control surfaces. For example, the horizontal stabilizer  152  may include one or more elevators configured to provide pitch control, and the vertical stabilizers  154  may each include a rudder configured to provide yaw control. In some embodiments, the empennage  150  includes only rudders or only an elevator, and in other embodiments the empennage  150  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  140 . 
     Pitch control in forward flight may be achieved by tilting the central rotor  140  forward or aft. Pitch control in forward flight may also be achieved by simultaneously tilting both proprotors  136   l,    136   r  higher or lower relative to the longitudinal or x-axis. For example, simultaneous upward tilting of the proprotors  136   l,    136   r  can produce a nose-up pitching moment, and simultaneous downward tilting of the proprotors  136   l,    136   r  can produce a nose-down forward pitching moment. 
     Roll control in forward flight may be achieved by tilting the central rotor  140  left or right. However, controlling roll by tilting the central rotor  140  requires a lateral tilting mechanism for the central rotor  140 . In some embodiments, the lower rotor hub  300  and/or the upper rotor hub assembly  400  may be simplified by providing only fore and aft tilting, and not lateral tilting. In such embodiments, aircraft roll can be achieved based on differential tilting of the proprotors  136   l,    136   r.  For example, tilting the left proprotor  136   l  slightly upward relative to horizontal and/or tilting the right proprotor  136   r  slightly downward relative to horizontal produces a right or clockwise rolling moment. Tilting the right proprotor  136   r  slightly upward relative to horizontal and/or tilting the left proprotor  136   l  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  136   l,    136   r.  For example, varying the relative speeds of the proprotors  136   l,    136   r  such that the left proprotor  136   l  turns at a higher rotational velocity than the right proprotor  136   r  produces a right yawing moment. Varying the relative speeds of the proprotors  136   l,    136   r  such that the right proprotor  136   r  turns at a higher rotational velocity than the left proprotor  136   l  produces a left yawing moment. 
     In various embodiments, 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 embodiments, 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  136   l  upward, increasing the rotational speed of the left proprotor  136   l,  tilting the right proprotor  136   r  downward, and/or decreasing the rotational speed of the right proprotor  136   r.    
     In addition to vertical and forward flight modes, the aircraft  100  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  100  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  100  configured for vertical flight. In this configuration, the proprotors  136   l,    136   r  are in the VTOL/hover position illustrated in  FIG.  6   , and the central rotor  140  may be turning under power while substantially parallel to the proprotors  136   l,    136   r.  To transition to forward flight, the proprotors  136   l,    136   r  simultaneously tilt forward into the forward flight position illustrated in  FIG.  6   . At approximately the same time, the central rotor  140  is tilted rearward (e.g., higher toward the front of the aircraft and lower toward the rear of the aircraft). If the central rotor  140  was powered during the preceding vertical or hovering flight, the rotor motor may continue to turn the central rotor  140  until the aircraft is established in forward flight. If the central rotor  140  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  140  to an appropriate rotational speed to provide lift for forward flight. Once the aircraft  100  is established in forward flight, the rotor motor may be deactivated as the relative airflow against the central rotor  140  becomes fast enough to cause autorotation of the central rotor  140 . 
     The aircraft  100  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  100  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  100  configured for horizontal flight. In this configuration, the proprotors  136   l,    136   r  are in the forward flight position illustrated in  FIG.  6   , and the central rotor  140  is turning by free autorotation as the aircraft travels in forward flight. To transition to vertical flight, the proprotors  136   l,    136   r  simultaneously tilt upward into the VTOL/hover position illustrated in  FIG.  6   . The central rotor  140  is tilted forward (e.g., to a level orientation substantially parallel to the proprotors  136   l,    136   r ). In some embodiments, the proprotors  136   l,    136   r  may be tilted beyond vertical, such as between 5 degrees and 30 degrees aft of vertical, if desired to slow the forward airspeed of the aircraft  100 . As the airspeed decreases and/or as the central rotor  140  tilts back to a level orientation, the autorotation of the central rotor  140  may decrease and/or stop. The rotor motor may be activated to turn the central rotor  140  to provide additional lift and/or stabilization to the aircraft  100  during the vertical flight phase. 
     Throughout the preceding disclosure, proprotors  136   l,    136   r  are described as being independently tiltable and controllable in order to achieve various flight control functions.  FIGS.  7 - 10    further illustrate example components configured to implement the proprotor control features described herein.  FIG.  7    depicts a medial portion of the lateral boom  132  of an example aircraft  100  including proprotor support and control components located thereon.  FIG.  8    is a detailed view of the proprotor support and control components of  FIG.  7   . As shown in  FIG.  7   , in some embodiments, the boom  132  may comprise a tubular medial structure, and the distal arms  134   l,    134   r  may be narrower tubular components coaxially mounted partially within the tubular medial structure. As shown in  FIGS.  7  and  8   , the aircraft  100  includes a gearbox  200 , which may be located within the proprotor tilt control housing  130  illustrated in  FIGS.  1 - 4   . As shown in  FIG.  8   , the gearbox  200  includes master gears  204  independently driven by proprotor tilt servos  202 . The master gears  204  are positioned so as to mesh with slave gears  206 , which are coaxial with and rotationally fixed to inner ends of the distal arms  134   l,    134   r.  Thus, the tilt servos  202  can tilt the proprotors  136   l,    136   r  by rotating the master gears  202 . 
     As shown in  FIGS.  9  and  10   , a proprotor motor  210  is disposed at an outer end of each distal arm  134   l,    134   r  of the boom  132 . Each proprotor motor  210  is rotationally fixed with respect to the corresponding distal arm  134   l,    134   r,  such that the rotational axis of the proprotor motor  210  is perpendicular to the lateral axis along the distal arm  134   l,    134   r . Thus, rotation of a distal arm  134   l,    134   r  about the lateral axis, under control of the proprotor tilt servos  202 , produces fore and aft tilting of the rotational axis of the proprotor motor  210 . Accordingly, each of the tilting operations described above with respect to the proprotors  136   l,    136   r  may be achieved by actuating the proprotor tilt servos  202  individually or simultaneously. 
       FIGS.  11 - 15    illustrate an example lower rotor hub  300  of an example aircraft such as the aircraft  100 . As described above, the lower rotor hub  300  serves as an attachment point for the upper rotor hub assembly  400  and the central rotor  140 . The lower rotor hub  300  is also configured to provide tilting and powered rotation of the central rotor  140 , as will now be described. The lower rotor hub  300  includes a rotor mount shaft  302  having one or more mounting pin holes  303  extending therethrough, a central rotor slave gear  304  meshed with a central rotor master gear  306 , a central rotor drive motor  308 , a tilt bearing  310 , and a central rotor tilt servo  312 . 
     The rotor mount shaft  302  serves as a mounting point for the central rotor  140 . The central rotor  140 , not shown in  FIGS.  11 - 15   , may be mounted to the aircraft  100  by coupling the upper rotor hub assembly  400  to the rotor mount shaft  302  and securing the upper rotor hub assembly  400  using one or more pins extending through mounting pin holes  303 . Pin-based mounting of the upper rotor hub assembly  400  allows the central rotor  140  to be easily and quickly attached to or detached from the aircraft  100  for transportation. 
     The central rotor slave gear  304  is coaxial with the rotor mount shaft  302  and is configured to transfer rotational motion of the central rotor master gear  306  to the rotor mount shaft  302  to drive the central rotor  140  during powered operation of the central rotor  140 . In some embodiments, the central rotor slave gear  304  is coupled to the rotor mount shaft  302  by a clutch mechanism and/or a one-way bearing (e.g., a one-way bearing  305 ) such that rotational motion of the central rotor slave gear  304  in a first direction (e.g., clockwise) is transferred to the rotor mount shaft  302 , but the rotor mount shaft  302  is free to spin in the same direction (e.g., clockwise) when the central rotor slave gear  304  is not rotating or is rotating more slowly than the rotor mount shaft  302 . Thus, the central rotor drive motor  308  can power the central rotor  140  (e.g., during vertical flight and/or during the transition from vertical flight to forward flight) by turning the central rotor master gear  306 , which in turn causes the central rotor slave gear  304  and rotor mount shaft  302  to rotate. 
     The central rotor tilt servo  312  is configured to tilt the lower rotor hub  300  relative to the mast  120 . Actuation of the central rotor tilt servo  312  causes rotation of the lower rotor hub  300  about the tilt bearing  310 , which accommodates motion about a lateral axis perpendicular to the rotor mount shaft  302 . 
       FIG.  14    illustrates an example forward-most tilt position of the lower rotor hub  300 . In the tilt position illustrated in  FIG.  14   , the rotor mount shaft  302  is substantially aligned with a vertical axis of the aircraft  100 , and a spinning central rotor  140  attached to the lower rotor mount  300  would produce a lifting force directly upward. The tilt position illustrated in  FIG.  14    may be used, for example, during vertical flight while the proprotors  136   l,    136   r  are in the VTOL/hover position illustrated in  FIG.  6   . 
       FIG.  15    illustrates an example rear-most tilt position of the lower rotor hub  300 . In the tilt position illustrated in  FIG.  15   , the rotor mount shaft  302  is tilted rearward approximately 20 degrees relative to vertical, such that a spinning central rotor  140  attached to the lower rotor mount  300  would rotate within a plane tilted approximately 20 degrees relative to a longitudinal axis of the aircraft  100 . The tilt position illustrated in  FIG.  15    may be used, for example, during horizontal flight while the proprotors  136   l,    136   r  are in the forward flight position illustrated in  FIG.  6   . The central rotor tilt servo  312  tilts the lower rotor hub  300  between the positions illustrated in  FIGS.  14  and  15    when the aircraft  100  transitions from vertical flight to forward flight, or from forward flight to vertical flight, as described above. 
       FIGS.  16 - 20    depict an example upper rotor hub assembly such as the upper rotor hub assembly  400  of the aircraft  100 .  FIG.  16    depicts the upper rotor hub assembly  400  including two mounting members  402 ,  404  coupleable to the lower rotor hub  300  of  FIGS.  11 - 15   , as well as an inner portion of a rotor blade  142  of the central rotor  140 .  FIGS.  17  and  18    depict the mounting members  402 ,  404  separately, including an example rotor blade  142  mounted to the first mounting member  402 .  FIGS.  19  and  20    depict the upper rotor hub assembly  400  in a folded configuration. 
     As described above, the central rotor  140  of the aircraft  100  functions similarly to the rotor of a gyroplane when the aircraft  100  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  100  is desired, it may be cumbersome to remove the blades  142  from the upper rotor hub assembly  400  for transport due to the time required to carefully align the blades  142  when reattaching them. As will now be described, the upper rotor hub assembly  400  is easily detachable from and attachable to the lower rotor hub  300 , and may be folded such that the entire central rotor  140  and upper rotor hub assembly  400  may be transported in a compact form while the blades  142  remain attached to the mounting members  402 ,  404 . 
     The upper rotor hub assembly  400  includes a first mounting member  402  and a second mounting member  404 . The mounting members  402 ,  404  each include mounting pin holes  405  configured to align with the mounting pin holes  303  of the lower rotor hub  300  of  FIGS.  11 - 15   . As shown in  FIG.  16   , the mounting members  402 ,  404  are shaped so as to nest together when both mounting members  402 ,  404  are mounted to the lower rotor hub  300 . In addition, the mounting members  402 ,  404  are shaped such that, when nested together in a mounted configuration, each mounting member  402 ,  404  can teeter independently about the axis of its mounting pin. In some embodiments, the aircraft  100  can be operated with only two blades  142  by using only the first mounting member  402  or only the second mounting member  404 . Two-bladed operation is associated with less lift and less induced drag during forward flight, and may be desirable when the aircraft  100  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  100  is carrying a relatively heavy load and greater lifting capacity is required. 
     Each mounting member  402 ,  404  includes two mounting brackets  406  disposed on opposite sides of the mounting member  402 ,  404  about the center of the hub. Each mounting bracket  406  includes two or more mounting holes  408  spaced apart to align with corresponding mounting holes  409  within the rotor blades  142 . In some embodiments, more than two mounting holes  408 ,  409 , such as three or more mounting holes, are provided in order to form a structurally robust connection between the rotor blades  142  and the mounting brackets  406 . 
     The mounting members  402 ,  404  each include two hinges  410  disposed between the mounting brackets  406  and the central portion of the mounting members  402 ,  404 . The hinges  410  allow the mounting brackets  406  and rotor blades  142  to be folded about the axis of the hinges  410  while the aircraft is not in flight. Each mounting bracket  406  is rigidly coupled to locking plates  414  having holes located to align with locking pin holes  413  of the central portion of the mounting members  402 ,  404 . When a hinge  410  is in the fully extended position (e.g., for flight), the hinge  410  may be locked in the fully extended position by inserting a locking pin  412  through locking pin holes  413  and the adjacent locking plates  414 . The locking pins  412  may include a retaining mechanism such as spring-loaded retaining balls or the like, to prevent the locking pins  412  from pulling out during operation. 
     In some embodiments, instead of or in addition to being foldable, the mounting members  402 ,  404  may also permit the rotor blades  142  to be removed from the mounting members  402 ,  402  in a manner that retains the alignment of the rotor blades  142  when they are reinserted. For example, the rotor blades  142  may be slidably mounted along one or more rails disposed within the mounting brackets  406 . Each mounting bracket  406  may include a release button which, when depressed, permits a rotor blade  142  within the mounting bracket  406  to slide outward to be removed. In some embodiments, the mounting brackets  406  and/or the mounting members  402 ,  404  may include a bayonet-type mounting system which maintains the appropriate alignment between opposing rotor blades  142 . Bayonet-type mounting systems will be described in greater detail with reference to  FIGS.  25  and  26   . 
       FIGS.  19  and  20    illustrate the folded configuration of mounting members  402 ,  404 . As shown in  FIGS.  19  and  20   , the removal of the locking pins  412  from the mounting members  402 ,  404  allows the locking plates  414  to move away from the locking pin holes  413 , allowing the mounting members  402 ,  404  to fold at the hinges  410  into a folded configuration. In the folded configuration, both mounting brackets  406  and rotor blades  142  attached to each mounting member  402 ,  404  are substantially parallel. Thus, a two-bladed or four-bladed central rotor  140  can be transported within a relatively narrow rectangular container having a length only slightly longer than each of the rotor blades  142 . This folding configuration is substantially more efficient than transporting the central rotor  140  in a flight configuration, which would require a container having two perpendicular dimensions of at least twice the length of each rotor blade  142 . Moreover, in some cases it may be desirable to fold the central rotor  140  while the central rotor  140  remains attached to the aircraft  100 . The folding mechanism described herein permits the two or four rotor blades  142  to be folded upward such that the horizontal footprint of the aircraft  100  may be minimized while it is being stored on the ground. 
       FIGS.  21  and  22    depict the upper rotor hub assembly  400  of  FIGS.  16 - 20    coupled to the lower rotor hub  300  of  FIGS.  11 - 15   . Each mounting member  402 ,  404  is coupled to the lower rotor hub  300  by a mounting pin  416  passing through the mounting pin holes  405  of the mounting member  402 ,  404  and the mounting pin holes  303  of the lower rotor hub  303 . Similar to the locking pins  412  that lock the mounting members  402 ,  404  in the fully extended position, the mounting pins  416  may include a retaining mechanism such as spring-loaded retaining balls or the like, to prevent the mounting pins  416  from pulling out during operation. 
       FIGS.  23  and  24    illustrate a further example configuration of an upper rotor hub assembly  400 . The example upper rotor hub assembly  400  of  FIGS.  23  and  24    further includes an upper rotor hub  416  which serves as a removable attachment point for the mounting members  402 ,  404 . The mounting members  402 ,  404  may be coupled to the upper rotor hub  416  such as by similar mounting pins  412 . In some embodiments, the mounting pins  412  coupling the mounting members  402 ,  404  to the upper rotor hub  416  may be permanent or semi-permanent pins, rather than the easily removable pins of  FIGS.  16 - 22   . 
     As shown in  FIG.  24   , the upper rotor hub  416  may in turn be connected to the rotor mount shaft  302  of the lower rotor hub by a further mounting pin  418 , which may similarly include spring-based or other retaining elements configured to be quickly releasable. Thus, the entire upper rotor mount hub assembly  400  may be folded and subsequently removed from the aircraft by removing a single mounting pin  418 . 
       FIGS.  25  and  26    illustrate a further example configuration of an upper rotor hub assembly  400  in which a bayonet-type mounting system allows the rotor blades  142  to be attached and detached from the upper rotor hub assembly  400 . Similar to the configuration of  FIGS.  23  and  24   , the upper rotor hub assembly  400  of  FIGS.  25  and  26    includes two mounting members  402 ,  404  attached to an upper rotor hub  416 . Each rotor blade  142  is mounted to a mounting bracket  406  by two, three, or more fasteners. Each mounting bracket  406  is fixed to a mounting body  420 , a portion or all of which may be integrally formed with the mounting bracket  406 . A blade attachment pin hole  422  extends laterally through the mounting body  420 . 
     Each opposing end of each mounting member  402 ,  404  includes a mounting body opening  424  sized and shaped to receive a mounting body  420 . Additional blade attachment pin holes  426  extend laterally through the sides of the mounting members  402 ,  404 . Thus, as illustrated by the transition from  FIG.  25    to  FIG.  26   , each rotor blade  142  may be mounted by sliding the mounting body  420  into a mounting body opening  424  until the blade attachment pin holes  422  are substantially aligned with blade attachment pin holes  426 . A blade attachment pin  428  may then be inserted through the blade attachment pin holes  422 ,  426  to secure the rotor blade  142  to the upper rotor hub assembly  400 . Dismounting of the blades, such as for storage, transport, etc., may be accomplished by removing the blade attachment pin  428  from the blade attachment pin holes  422 ,  426  and subsequently sliding the mounting body  420  out of the mounting body opening  424 . 
     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.  27    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.  FIGS.  28  and  29    illustrate side and cross-sectional views of the rotor hub assembly of  FIG.  27    in a zero pitch configuration.  FIGS.  30  and  31    illustrate side and cross-sectional views of the rotor hub assembly of  FIGS.  27 - 29    in a positive pitch configuration.  FIG.  32    depicts an exploded view of the rotor hub assembly of  FIGS.  27 - 31   . The rotor assembly illustrated in  FIGS.  27 - 32    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.  FIGS.  28  and  29    illustrate the assembly with the rotor blades being at a zero-degree pitch, and  FIGS.  30  and  31    illustrate the assembly with the rotor blades being at a higher pitch that creates lift. 
     As shown in  FIGS.  28 ,  30 , and  32   , 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  FIGS.  25  and  26   , 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  1  and  2  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  FIGS.  27 - 32    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  1  and  2  slide within the rotor orientation tracks, allowing the gyrocopter to fly forward with the aid of the pusher motor/propeller assembly or, in any of the various embodiments including two laterally mounted tiltable or non-tiltable proprotors (e.g., proprotors  136   l,    136   r ), allowing the gyrocopter to fly forward with the aid of the left and right proprotors. 
     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 250-450 RPM, does not create sufficient centrifugal force to collapse the gas spring outward, while a higher RPM range, such as above 550 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 450 RPM or slower while in horizontal flight, and with the main rotor turning at 550 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 450 RPM and 550 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.  33    illustrates an example gearing system for increasing the RPM of the upper rotor hub assembly of  FIGS.  27 - 32   .  FIG.  34    is a cross-sectional view illustrating a braking system for reducing the rotational speed of the upper rotor hub assembly of  FIGS.  27 - 32   . 
     As shown in  FIG.  33   , 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  FIGS.  1 - 4   ) 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, at a full stop, or in a reversed direction, in order to fully counter the torque of the rotor head assembly. Thus, the two proprotors can be independently controlled to operate at different rotational velocities and/or in different rotational directions. In some embodiments, the other proprotor (e.g., the right proprotor) may even be rotated 180 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.  34   , 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. 
       FIGS.  35 - 37    illustrate an example aircraft including a tiltable empennage. In some embodiments, the empennage, including the horizontal stabilizer and one or two rudders, are turned 90 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).  FIGS.  35  and  36    illustrate the empennage in an upright position, such as for forward flight.  FIG.  37    illustrates the empennage in a lowered position, tilted approximately 90 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 90 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. 
       FIGS.  38 - 41    illustrate a further example embodiment of an aerial vehicle including a single proprotor and pivotable wings configured to operate as control surfaces. The aerial vehicle of  FIGS.  38 - 41    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  FIGS.  38 - 41    may be implemented with any of the aerial vehicle embodiments disclosed herein. 
     As shown in  FIGS.  38  and  39   , in horizontal forward flight, the two wings have substantially the same horizontal orientation such that the wings are streamlined for forward flight. In some embodiments, 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  FIGS.  40  and  41   , 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 90 degrees and the right wing is pivoted by less than 90 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  FIGS.  40  and  41    creates a counterclockwise torque that counters the clockwise torque created by the powered rotor. 
       FIGS.  42  and  43    illustrate a further example configuration of an aircraft  100  in accordance with the technology described herein. As shown in  FIGS.  42  and  43   , the aircraft  100  includes side-mounted proprotors  136   l,    136   r  in a fixed (e.g., non-tiltable) configuration. Empennage  150  is fixed relative to the aircraft  100  and can be free of control surfaces such as rudders or elevators. A tilt bearing  310  is configured to allow fore-aft tiling, lateral tilting, and/or a combination of fore-aft and lateral tilting of the central rotor. Motors driving the proprotors  136   l,    136   r  may be configured to individually and/or differentially control the speed and/or rotational direction of proprotors  136   l,    136   r.  For example, to create a yawing moment to counter the torque created by a clockwise-rotating central rotor, the left proprotor  136   l  may be powered to rotate in a first direction at a desired rotational speed to produce forward thrust while the right proprotor  136   r  is powered to rotate in a second direction opposite the first direction (e.g., a reversed rotational direction) at a desired rotational speed to produce reverse thrust, or is powered to rotate in the first direction at a lower rotational speed relative to the rotational speed of the left proprotor  136   l,  such that a rightward yawing moment is created. An opposite counter-torque scheme may be implemented for a counterclockwise-rotating central rotor. The variable rotational speeds and/or directions of the proprotors  136   l,    136   r  may be controlled in coordination with fore-aft and/or lateral tilting of the central rotor at the tilt bearing  310  (e.g., under control of an autopilot or other control circuitry) to selectively control roll, pitch, yaw, and thrust of the aircraft during forward flight, as described elsewhere herein. 
       FIGS.  44  and  45    illustrate an example rotor assembly  440  in accordance with the present technology. An upper rotor hub assembly includes a central section  442  and mounting brackets  444  configured as a bayonet-type blade mounting system and configured to receive rotor blades therein, as described herein with reference to  FIGS.  25  and  26   . A blade pitch adjustment system  446  provides for automatic adjustment of rotor blade pitch based on rotational velocity, as described herein with reference to  FIGS.  27 - 32   . Rotor tilt servos  448 ,  450 ,  452 , and  454  are provided to control fore-aft and/or lateral tilting of the upper rotor hub assembly. Thus, a motor  456  or corresponding lower rotor hub assembly may remain fixed relative to a fuselage of an aircraft (e.g., aircraft  100  as described elsewhere herein) in which the rotor assembly  440  is installed, while rotor tilt servos  448 ,  450 ,  452 ,  454  selectively tilt the upper rotor hub assembly and connected rotor blades to control pitch and/or roll of the aircraft. In some embodiments, the rotor assembly  440  may include only two tilt servos, for example, servos  448  and  450 , with one of servos  448  and  450  oriented to control fore-aft tilting, and the other one of servos  448  and  450  oriented to control lateral tilting. Additional tilt servos  452 ,  454  may be included for redundancy and/or for additional stability when controlling tilt of the upper rotor hub assembly. 
     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. Furthermore, various omissions, substitutions, and changes in the systems and methods described herein may be made without departing from the spirit of the disclosure. The accompanying claims and their equivalents are intended to cover such forms or modifications as would fall within the scope and spirit of the disclosure. Accordingly, the scope of the present disclosure 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. All of the features disclosed in this specification (including any accompanying claims, abstract and drawings), and/or all of the steps of any method or process so disclosed, may be combined in any combination, except combinations where at least some of such features and/or steps are mutually exclusive. The protection is not restricted to the details of any foregoing embodiments. The protection extends to any novel one, or any novel combination, of the features disclosed in this specification (including any accompanying claims, abstract and drawings), or to any novel one, or any novel combination, of the steps of any method or process so disclosed. 
     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. Conversely, various features that are described in the context of a single implementation can also be implemented in multiple implementations separately or in any suitable subcombination. Moreover, although features may be described above as acting in certain combinations, one or more features from a claimed combination can, in some cases, be excised from the combination, and the combination may be claimed as a subcombination or variation of a subcombination. 
     Moreover, while operations may be depicted in the drawings or described in the specification in a particular order, such operations need not be performed in the particular order shown or in sequential order, or that all operations be performed, to achieve desirable results. Other operations that are not depicted or described can be incorporated in the example methods and processes. For example, one or more additional operations can be performed before, after, simultaneously, or between any of the described operations. Further, the operations may be rearranged or reordered in other implementations. Those skilled in the art will appreciate that in some embodiments, the actual steps taken in the processes illustrated and/or disclosed may differ from those shown in the figures. Depending on the embodiment, certain of the steps described above may be removed, others may be added. Furthermore, the features and attributes of the specific embodiments disclosed above may be combined in different ways to form additional embodiments, all of which fall within the scope of the present disclosure. Also, the separation of various system components in the implementations described above should not be understood as requiring such separation in all implementations, and it should be understood that the described components and systems can generally be integrated together in a single product or packaged into multiple products. 
     For purposes of this disclosure, certain aspects, advantages, and novel features are described herein. Not necessarily all such advantages may be achieved in accordance with any particular embodiment. Thus, for example, those skilled in the art will recognize that the disclosure may be embodied or carried out in a manner that achieves one advantage or a group of advantages as taught herein without necessarily achieving other advantages as may be taught or suggested herein. 
     Conditional language, such as “can,” “could,” “might,” or “may,” unless specifically stated otherwise, or otherwise understood within the context as used, is generally intended to convey that certain embodiments include, while other embodiments do not include, certain features, elements, and/or steps. Thus, such conditional language is not generally intended to imply that features, elements, and/or steps are in any way required for one or more embodiments or that one or more embodiments necessarily include logic for deciding, with or without user input or prompting, whether these features, elements, and/or steps are included or are to be performed in any particular embodiment. 
     Conjunctive language such as the phrase “at least one of X, Y, and Z,” unless specifically stated otherwise, is otherwise understood with the context as used in general to convey that an item, term, etc. may be either X, Y, or Z. Thus, such conjunctive language is not generally intended to imply that certain embodiments require the presence of at least one of X, at least one of Y, and at least one of Z. 
     The scope of the present disclosure is not intended to be limited by the specific disclosures of preferred embodiments in this section or elsewhere in this specification, and may be defined by claims as presented in this section or elsewhere in this specification or as presented in the future. The language of the claims is to be interpreted broadly based on the language employed in the claims and not limited to the examples described in the present specification or during the prosecution of the application, which examples are to be construed as non-exclusive.