Abstract:
A rotorcraft having two coaxial, counter-rotating rotors, one proximate to the forward end of the fuselage and one proximate to the aft end of the fuselage, that generate the forces necessary to lift the craft and maneuver it in the air by adjusting the pitch of the rotor blades throughout their rotation, and a method of flying a dual rotor rotorcraft involving taking off in a vertical orientation, climbing vertically, transitioning to generally horizontal flight, flying horizontally, and subsequently repeating the sequence in reverse to land again in a vertical orientation.

Description:
BACKGROUND 
     The present invention relates to an aircraft structure wherein conventional wings are omitted and two, coaxial, counter-rotating, rotors utilize high speed cyclic pitch and collective pitch control to generate the forces necessary to lift the craft, maneuver it, and propel it through the air. A method of flight utilizing such a craft is also disclosed. 
     Dual coaxial rotor rotorcraft are known in the art. However, such rotorcraft typically cannot operate in a generally horizontal orientation without utilizing fixed wings to generate lift. Because fixed wings contribute to skin friction drag at cruise speeds, an air vehicle, capable of generating the necessary lift and directional forces for vertical take off and landing, vertical hovering, and horizontal flight, without the use of fixed wings is desirable. 
     Accordingly, it is an object of the present invention to provide a dual coaxial rotor rotorcraft capable of taking off vertically, hovering, transitioning to horizontal flight, returning to vertical hovering, and landing vertically, all without the benefit of fixed aerodynamic appendages. 
     SUMMARY 
     According to the present invention, there is provided a dual coaxial rotor rotorcraft and a method of flight. The rotorcraft includes a fuselage having a forward end and an aft end, and two co-axial, counter-rotating rotor assemblies, one of which is proximate to the forward end of the fuselage and one of which is proximate to the aft end of the fuselage. Each rotor assembly further includes high-speed actuators capable of adjusting the pitch of each blade independently and continuously throughout the entire rotational cycle of each blade. The actuators are controlled wirelessly by an onboard control system having a computer-based controller, at least one accelerometer, at least one rate gyroscope, a wireless interface, and a wireless transceiver. The onboard control system wirelessly receives flight commands from a remote flight control module through the wireless transceiver, determines the rotor speed and pitch angles necessary at each rotational point, and directs the rotor drive motors to adjust the speed of the rotors and wirelessly directs the high speed actuators through the wireless interface to adjust the pitch of each blade such that the rotors generate the necessary forces and the vehicle executes the flight commands. 
     The method of flight is directed to a dual coaxial rotor rotorcraft having one rotor assembly proximate to the forward end of the fuselage and one rotor assembly proximate to the aft end of the fuselage, each such rotor assembly including two or more independently controllable, continuously-variable-pitch blade assemblies. The method includes the steps of taking off from an orientation in which the rotor axis is generally vertical, adjusting the pitch of the rotor blades to generate forces perpendicular to the rotorcraft axis to cause the craft to pitch forward in the direction of travel until the rotorcraft is in a generally horizontal orientation, flying in a generally horizontal orientation, transitioning back to a generally vertical orientation by adjusting the pitch of the rotor blades to generate forces perpendicular to the rotorcraft axis to cause the craft to pitch nose up until the rotorcraft is in a generally vertical orientation, and landing in a generally vertical orientation. 
     Other features and advantages of the invention will become apparent from the following description, including the drawings and the claims. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       Other features in the invention will become apparent from the attached drawings, which illustrate certain preferred embodiments of the rotorcraft and method of this invention, wherein 
         FIG. 1  is a perspective view of one preferred embodiment of the rotorcraft of the present invention; 
         FIG. 2  is a side view, partially in section, of one preferred embodiment of the rotorcraft of the present invention having dual rotor drive motors and showing internal components in block diagram form; 
         FIG. 3  is a side view, partially in section, of one preferred embodiment of the rotorcraft of the present invention having a single rotor drive motor and showing internal components in block diagram form; 
         FIG. 4  is a side view of one preferred embodiment of the rotorcraft of the present invention illustrating internal wireless communications components in phantom; 
         FIG. 5  is a perspective view of a rotor assembly of one preferred embodiment of the rotorcraft of the present invention; 
         FIG. 6  is a perspective view of a rotor drive gear assembly and rotor drive motor of one preferred embodiment of the rotorcraft of the present invention; 
         FIG. 7  is an exploded perspective view of the rotor drive and rotor assemblies of one preferred embodiment of the rotorcraft of the present invention; 
         FIG. 8  is an exploded perspective view of the rotor drive and rotor assemblies of one preferred embodiment of the rotorcraft of the present invention; 
         FIG. 9  is an exploded, perspective view of the fuselage components of one preferred embodiment of the rotorcraft of the present invention; 
         FIG. 10  is a plan view of the remote flight control module of one preferred embodiment of the rotorcraft of the present invention, 
         FIG. 11  is a perspective view of one preferred embodiment of the rotorcraft of the present invention with landing gear, 
         FIG. 12  is a partially exploded, perspective view of the deployed landing gear of one preferred embodiment of the rotorcraft of the present invention, 
         FIG. 13  is a fragmentary, perspective view of the retracted landing gear of one preferred embodiment of the rotorcraft of the present invention, with portions of the rotorcraft shown in phantom, 
         FIGS. 14-A  and  14 -B are a schematic view of the flight method of the present invention, 
         FIG. 15  is a perspective view of one preferred embodiment of the rotorcraft of the present invention illustrating lift vectors generated in vertical hover, 
         FIG. 16  is a perspective view of one preferred embodiment of the rotorcraft of the present invention illustrating lift vectors generated in horizontal flight, 
         FIG. 17 . is a perspective view of one preferred embodiment of the rotorcraft of the present invention illustrating lift vectors generated in the transition from vertical flight to horizontal flight, and 
         FIG. 18  is a perspective view of one preferred embodiment of the rotorcraft of the present invention illustrating lift vectors generated when maneuvering in a vertical orientation. 
     
    
    
     DETAILED DESCRIPTION 
     While the following describes preferred embodiments of the rotorcraft and method of this invention, it is to be understood that this description is to be considered only as illustrative of the principles of the invention and is not to be limitative thereof, as numerous other variations, all within the scope of the invention, will readily occur to others. In this specification, the term “adapted” shall mean configured, dimensioned, oriented and arranged as appropriate. 
       FIGS. 1-18  illustrate preferred embodiments of the rotorcraft and methods of this invention. The rotorcraft  1 , which may conveniently be a remotely piloted vehicle, generally comprises a fuselage  2  having a forward end  4  and a rear, or aft, end  6 , with a longitudinal axis  8  extending generally through said forward end  4  and said aft end  6 . The rotorcraft  1  further comprises two counter rotating rotor assemblies  10  ( FIG. 5 ), one of which is a forward rotor assembly  110  proximate to the forward end  4  and one of which is an aft rotor assembly  210  proximate to the aft end  6 , and each of which is coaxial with longitudinal axis  8 . The fuselage  2  is preferably in the form of a cylinder with tapering ends  4 ,  6 , but may include bulges (not illustrated) as necessary to accommodate internal components and payloads. The rotor assemblies  10  are driven by a rotor drive system  14  as shown in  FIG. 3 . The rotor drive system  14  comprises at least one rotor drive motor  16  mechanically connected to at least one rotor drive gear set shown in  FIG. 6 , said gear set(s) being indicated as elements  118  and  218  in  FIGS. 2 and 3  and being mechanically connected, respectively, to forward rotor assembly  110  and aft rotor assembly  210 , and geared so that rotor assemblies  110 ,  210  are kept rotating in opposite directions at substantially the same speed, thereby helping avoid instability and unwanted rotation of fuselage  2 . As shown in the illustrations, the preferred embodiment of the rotorcraft has no wings or attached airfoils, but only the rotating rotor assemblies for aerodynamic for lift and control. 
     As shown in  FIGS. 7 and 8 , each rotor assembly  10  comprises a rotor battery pack  20 , preferably being made up of five Kokam rechargeable lithium-polymer cells, wired in series, per variable pitch blade assembly, a rotor blade controller  22  electrically connected to rotor battery pack  20 , wireless rotor blade interface  24  electrically connected to rotor blade controller  22 , and at least two, but preferably four, continuously variable pitch blade assemblies. Preferably, all four rotor battery packs  20  are connected and jointly provide electrical power to wireless rotor blade interface  24  and rotor blade controller  22 , in addition to actuator drive amplifier  36  and actuator drive  32  discussed more fully below. Wireless rotor blade interface  24  wirelessly receives commands from the onboard controller  25 , which is illustrated in  FIGS. 2 and 3  and is discussed below, and relays those commands to rotor blade controller  22 . Rotor blade controller  22  then directs the blade assemblies, discussed more fully below, to adjust the pitch of blades  28  at each point in their rotation about the fuselage  2 . Because rotor battery pack  20  and rotor blade controller  22  are physically separate from the onboard controller  25 , there is no need for mechanical or electrical linkages between onboard controller  25  and the rotating blade assemblies. Instead, the rotor assemblies  10  are self-contained units and onboard controller  25  utilizes wireless communications to direct rotor blade controllers  22  to continuously adjust the pitch of each of each blade  28  as it rotates about fuselage  2 . 
     There is preferably one rotor blade assembly for each blade  28  on each rotor assembly  10 . The rotor blade assemblies shown in  FIGS. 5 ,  7 , and  8  comprise a high speed electric actuator drive  32 , such as the brushless DC motors available from ARC Systems, electrically connected to rotor blade controller  22 , actuator drive amplifier  36 , and rotor battery pack  20 . Actuator drive amplifier  36 , which may conveniently be a brushless motor servo amplifier such as those available from Advanced Motion Controls, accepts torque commands from rotor blade controller  22  and causes actuator drive  32  to produce the commanded torque. Actuator drive  32  is mechanically connected to rotor blade  28 , preferably through an actuator gear  34 . Upon receiving signals from onboard controller  25 , rotor blade controller  22  causes actuator drive amplifier  36  and high speed actuator drive  32  to adjust the pitch, or angle of attack, of each rotor blade  28 . 
     Preferably, each rotor blade  28  is a low profile propeller blade of the type commonly used on tilt rotor aircraft and well known to those of ordinary skill in the art. Blade pitch actuator drive  32  and actuator gear  34  are preferably adapted to continuously adjust the pitch angle of the rotor blade through at least 120 degrees. For successful flight operations, it is important that wireless interface  24 , rotor blade controller  22 , actuator drive amplifier  36 , actuator drive  32 , and actuator gear  34  be adapted to operate at a speed sufficient to be able to finely adjust the pitch of each rotor blade  28  throughout its rotation. As is described in more detail below, the pitch angle of each blade  28  may vary as much as 120 degrees or more in the course of a single rotation of the blade  28  about axis  8 . A low backlash in actuator gear  34  is important both in obtaining precise orientations and fast positioning. In one preferred embodiment, precision spur gears, such as those available from Stock Drive Products may suitably be utilized with an actuator drive gear ratio of 32:9. As will be clear to those of ordinary skill in the art, the required high speed internal wireless communication can be achieved through the use a wideband radio-frequency (RF) data link  30  shown in  FIG. 4 , preferably a wireless high speed digital data network such as IEEE 802.11 or others well known to those of ordinary skill in the art, and protocol (not illustrated) between onboard controller  25  and rotor blade controllers  22 , in combination with a high speed central processing unit (CPU) (not illustrated) in rotor blade controllers  22  programmed to communicate with onboard controller  25 . In addition to receiving commands from onboard controller  25 , the CPU may conveniently receive feedback data from conventional low profile optical encoders (not illustrated), such as those available from Renco and capable of encoding 2048 lines per revolution at 8000 RPM. Such encoders may conveniently be electrically connected to rotor blade controller  22  and adapted to read the rotational position of conventional sensor rings (not illustrated) mechanically connected to the shafts of actuator drives  32  whereby the pitch angle of each blade may be determined by rotor controller  22 , and wirelessly communicated to onboard controller  25  via data link  30 . Based on the information from the encoders and the commands from onboard controller  25 , the rotor blade controller  22  generates commands to the actuator drive amplifier  36 , which in turn powers actuator drive  32  to adjust the pitch of the rotor blades  28  throughout the blade&#39;s rotation about axis  8 . 
     An onboard controller  25  ( FIGS. 2-3 ) is responsible for receiving flight commands, determining how the speed of rotor assemblies  10  and pitch of the individual blades  28  need to be adjusted for the vehicle  1  to execute the flight commands, and directing the rotor drive motor  16  and rotor blade controllers  22  to make the necessary adjustments in rotor speed and blade pitch. Onboard controller  25  comprises a computer-based controller  38 , at least one three axis accelerometer  40 , at least one three-axis rate gyroscope  42 , an onboard wireless interface  44  adapted to communicate with the rotor assembly wireless interfaces  24 , and an onboard transceiver  46  and antenna  48  adapted to wirelessly communicate with a remote flight control module  50  ( FIG. 10 ), described more fully below. The computer-based controller  38  comprises a high speed CPU (not illustrated) such as the P501 from General Microsystems, programmed to translate flight commands received from the remote flight control module  50  into commands to adjust rotor speed and blade pitch such that the rotor blades generate the lift, rotational, and directional forces necessary for rotorcraft  1  to execute flight commands. The onboard controller  25  and rotor drive system  14  are powered by a main battery pack  52 , preferably comprising rechargeable lithium-polymer batteries of the type that are well known to those of ordinary skill in the art. The computer-based controller  38  is electrically connected to three-axis accelerometer  40  and three-axis rate gyroscope  42 , both of which may preferably be powered by main battery pack  52 , and both of which provide real time data indicating angular rates and acceleration to the CPU of the computer-based controller  38  and enabling computation of orientation and velocity. Computer-based controller  38  also receives flight commands from remote flight control module  50  via another RF link. As is discussed in more detail below, the RF link to remote flight control module  50  is bi-directional, thereby allowing computer-based controller  38  to communicate flight and related data back to remote flight control module  50 . The programming in the computer-based controller  38  analyzes the flight commands received from the remote flight control module  50  together with the data from accelerometer  40  and rate gyro  42  and the rotor speed and pitch of each blade  28  at each point in rotation, and calculates the changes required in the speed of each rotor and the pitch of each blade in order to execute the flight command. Computer based controller  38  then directs rotor drive system  14  and movement of the blade assemblies as required. 
     Remote flight control module  50  is a remote station from which rotorcraft  1  may be flown. Remote flight control module  50  comprises a set of manually operable flight controls  53  and a wireless flight control transceiver  54  and antenna  56  adapted to wirelessly communicate with the onboard transceiver  46 . It is also may comprise a display  58  capable of displaying flight, operational, system status, and related data to the operator. Remote flight control module  50  may be in the form of a sit-down flight operations console (not illustrated), or a handheld unit  50 . In the handheld unit embodiment, flight controls  53  may conveniently comprise a first joy stick  60  adapted such that fore-aft movement commands vehicle engine power and collective pitch and left-right movement commands left or right vehicle rotation and a second joystick  62  adapted such that fore-aft movement commands vehicle pitch and left-right movement commands vehicle roll. A landing gear control  64  commands deployment or retraction of landing gear  66  discussed more fully below. 
     For take off and landing, the rotorcraft of the present invention may include a landing gear  66  proximate to the aft end  6  of fuselage  2  and illustrated in  FIG. 11 . As is discussed above and illustrated in  FIG. 1 , the rotorcraft  1  of the present invention has a longitudinal axis  8  running from the forward end  4  through to the aft end  6  of fuselage  2 . Within this specification, rotorcraft orientation is discussed with respect to longitudinal axis  8  and the ground. Accordingly, said landing gear  66  is conveniently adapted to support rotorcraft  1  in a generally vertical orientation. Landing gear  66  comprises at least three struts  68  extending from fuselage  2  proximate to aft end  6 , and below aft rotor  210 , preferably in a tripod configuration. 
     For superior aerodynamic performance, landing gear  66  may be retractable. In such an embodiment, landing gear  66  may comprise at least 3, and preferably exactly 3, retractable struts  68 , and a electric landing gear motor  70 . Landing gear motor  70  is preferably powered by main battery pack  52 . To produce the commanded currents and resulting torques in landing gear motor  70 , landing gear motor amplifier  72  is utilized. Accordingly, computer based controller  38  is electrically connected to landing gear motor amplifier  72  which, in turn, is connected to landing gear motor  70 , said connections being adapted to enable onboard controller  25  to cause landing gear motor  70  to rotate in either direction, whereby landing gear struts  68  are deployed and retracted. 
     One embodiment of the structure of the retractable landing gear  66  of the present invention is illustrated in  FIGS. 11-13 . In such an embodiment, landing struts  68  comprise two telescoping sections. The uppermost section  74  being connected to a bulkhead  78  via a hinged mechanical connection  80 , and the lowermost section  76  being adapted to retract within uppermost section  74 . Uppermost section  74  may be movably attached to a support arm  82  running from said strut  68  to threaded runner  84  on a central threaded rod  86  that is, in turn, driven by landing gear motor  70  such that, upon rotation of threaded rod  86  in a first direction, runner  84  is driven down threaded rod  86 , thereby causing lowermost section  76  to retract and, upon rotation of threaded rod  86  in the opposite direction, said runner  84  is driven up threaded rod  86 , thereby causing landing strut  68  to deploy. A cable or line  88 , conveniently made of high strength monofilament line or other materials well known to those of ordinary skill in the art, may conveniently attach to arm  82  proximate to the connection to said runner  84  on one end, proceed through a pulley  90  or similar turning structure on strut  68 , and to the upper end of lowermost section  76  of strut  68 , and be adapted such that upon deployment, said line urges lowermost section  76  to extend. Internal to strut  68  there is preferably an extension spring (not illustrated) running from lowermost section  76  to a point proximate to the upper end of uppermost section  74 . The spring is adapted to keep tension on line  88  and to pull lowermost section  76  into uppermost section  74  as landing gear  66  retracts. 
     For aerodynamic efficiency, fuselage  2  may further comprise cowl sections  93  proximate to aft end  6 . Cowl sections  93  may conveniently be mechanically attached to uppermost section  74  and adapted to close, thereby covering landing gear  66  when retracted. Cowl sections  93  thereby open as retractable landing gear  66  deploys and close as landing gear  66  retracts. 
     As is discussed above, the rotor drive system  14  of the rotorcraft may conveniently comprise a single rotor drive motor  16  ( FIG. 3 ) or two sets of rotor drive motors ( FIG. 2 ), one set being a forward rotor drive motor set  162  and one being an aft rotor drive motor set  164 . In embodiments utilizing a single rotor drive motor  16 , motor  16  is mechanically connected to both the forward  110  and aft  210  rotor assemblies through forward gear set  118  and aft gear set  218  respectively. In such embodiments, aft gear set  218  is adapted to reverse the rotation of aft rotor assembly  210 , but otherwise maintain a generally equal rotational speed between forward rotor assembly  110  and aft rotor assembly  210 . Rotor drive motor  16  is mechanically connected to both forward gear set  118  and aft gear set  218 , and is electrically connected to computer based controller  38 . Drive motor  16  is also electrically connected to main battery pack  52 , preferably through drive motor amplifier  92 . Computer based controller  38  outputs a low level voltage command to drive motor amplifier  92  which then drives motor  16  with power from the main battery pack  52 . Drive motor  16  may conveniently be a brushless, direct current electric motor such as are well known to those of ordinary skill in the art. 
     In an alternate embodiment shown in  FIG. 2 , rotor drive system  14  may conveniently comprise two sets of rotor drive motors, one set being a forward rotor drive motor set  162  and one being an aft rotor drive motor set  164 . In one preferred embodiment, motors capable of generating 1200 W-1500 W continuously, such as the 1515/1.5 Y DC motors available from Neumotors, may be utilized to deliver the power necessary to drive rotor assemblies  10 . To power the motors, power amplifiers  92  rated at 125 A peak and 100 A continuous current, such as the SH-96 available from Aveox, are paired with each drive motor. Forward rotor drive motors  162  are mechanically connected to forward gear set  118  and aft rotor drive motors are mechanically connected to aft gear set  218 . Each such gear set  118  and  218  ( FIG. 6 ) may comprise a motor shaft gear  94  being centrally positioned between two equally-sized primary reduction gears  96 . Each primary reduction gear  96  may be mechanically connected to pinion gear  98  which, in turn, is adapted to engage a main drive gear  100  mounted to drive shaft  102 . Drive shaft  102  is preferably hollow, thereby allowing a space for electrical connections between computer based controller  38 , main battery pack  52 , rotor drive motors  160 , and landing gear drive motor  70 . As is shown in  FIGS. 6-8 , drive shaft  102  may also have attached to it magnetic ring  104  which is adapted to be read by encoder  106 , whereby computer-based controller  38  may determine the rotational position and speed of forward rotor assembly  110  and aft rotor assembly  210 , respectively. Utilizing the above-described gear set structure allows for multiple, and preferably four, drive motors  162  or  164  to simultaneously drive each rotor assembly  10 . In this configuration, and as is illustrated in  FIG. 8 , main battery pack  52  may comprise three pluralities of lithium polymer cells, one plurality  152  directly surrounding forward drive motor set  162 , one plurality  352  directly surrounding aft drive motor set  164  and a third plurality  252  positioned therebetween. In this way, main battery pack  52  may comprise a large number of individual cells as required to meet the power requirements of onboard controller  25 , rotor drive system  14  and landing gear  66 . Suitable cells for main battery pack  52  may be obtained from a variety of sources and wired in a variety of configurations that will be obvious to those of skill in the art. One such suitable configuration utilizes cells rated at 3.7V and 2000 mA-hr, such as those from Kokam. Such cells are capable of delivering a continuous 30 A current with a peak current of 60 A. Using such cells, pluralities  152  and  352  may comprise eight cells per drive motor  160 , for a total of 32 cells, wired to generate 29.6V per motor. Similarly center plurality  252  may conveniently comprise another group of 32 such cells, thereby providing a total of 96 main power cells. In such configurations, the cells may be wired as groups of 8 connected in series. 
     In addition to the rotorcraft  1  of present invention, a method of flying a dual rotor, vertical take off and landing rotorcraft  501 , which may suitably be substantially similar to rotorcraft  1 , is also disclosed and is illustrated in  FIGS. 14-18 . The rotorcraft  501  comprises a fuselage  502  having a top side or upwardly facing portion  514  when in horizontal flight, bottom side or downwardly facing portion  516 , a forward end  504 , and an aft end  506 , and two co-axial, counter-rotating rotor assemblies  511 ,  512  rotating about a longitudinal axis  508  extending generally through said forward end  504  and said aft end  506 . The rotor assemblies  511 ,  512  consist of forward rotor assembly  511  positioned proximate to forward fuselage end  504  and an aft rotor assembly  512  positioned proximate to aft fuselage end  506 . Each rotor assembly  511 ,  512  comprises two or more, but preferably four, independently controllable, continuously-variable-pitch blades  528 . In this specification, angles of rotation of rotor blades  528  about longitudinal axis  508  are discussed with reference to a point  515  on the top fuselage side  514  and the direction of rotation. Positive rotation of a rotor blade is taken to indicate an increasing rotation angle even though the rotors rotate in opposite directions. This point  515  can be selected at any point on upwardly facing portion, or top side  514  as shown on  FIGS. 14 and 16 . More particularly, a blade  528  that is pointed directly at point  515  ( FIG. 16 ) on the top fuselage side  514  has a rotation angle of zero. That blade  528  will pass through blade angles of 45 degrees, to 90 degrees, to 180 degrees, whereupon it is pointed directly at bottom fuselage side  516 . Similarly, pitch angle in this specification is discussed relative to the plane of rotation of rotor assemblies  511 ,  512 . Accordingly, a blade  528  that is aligned parallel with the plane of rotation is said to have a zero pitch angle. A blade  528  that is perpendicular to the plane of rotation is said to have a 90 degree pitch angle if the forward edge of the blade  528  is directed toward the forward fuselage end  504 , and a −90 degree pitch angle if the forward edge of blade  528  is directed toward aft fuselage end  506 . Blade angles required for flight in various orientations will vary based on the blade cross section, profile, and total blade area. 
     The power to weight ratio of rotorcraft  501  should preferably be between five and ten pounds per installed horsepower. Because the rotors preferably turn moderately fast the areas of blades  528  can preferably be made small. To provide ample margins of safety against blade stall, and because the blades may conveniently be thin with small camber, lift coefficients less than 0.8 may be used. The result is a rotorcraft  501  with total installed power requirements similar to that of a traditional helicopter, reasonable transition power requirements, and small total rotor blade surface area. 
     The method comprises the step of taking off from an orientation in which longitudinal axis  508  is generally vertical as is shown at the left side of  FIG. 14-A . Take off is accomplished by adjusting the pitch of each blade  528  to approximately equivalent angles of attack, thereby generating vertical lift without creating undesired horizontal forces. By maintaining the same torque on the two counter-rotating rotors  511 ,  512 , the torque generated by the rotors  511 ,  512  cancel, thereby preventing undesired rotation about longitudinal axis  508 . Typical blade angles of the blades  528  of forward rotor assembly  511  and aft rotor assembly  512  during vertical hover are summarized in the following table. Take off is accomplished by increasing the blade angles and rotational speed as required depending on payload. 
     
       
         
               
             
               
               
             
               
               
               
               
               
               
               
               
               
             
               
               
               
               
               
               
               
               
               
             
           
               
                   
               
               
                 Vertical Orientation: Hovering, craft pitch = 0° 
               
             
          
           
               
                   
                 Rotation angle of blade 
               
             
          
           
               
                   
                 0 
                 45 
                 90 
                 135 
                 180 
                 225 
                 270 
                 315 
               
               
                   
                   
               
             
          
           
               
                 FWD rotor blade angle 
                 15 
                 15 
                 15 
                 15 
                 15 
                 15 
                 15 
                 15 
               
               
                 AFT rotor blade angle 
                 15 
                 15 
                 15 
                 15 
                 15 
                 15 
                 15 
                 15 
               
               
                   
               
             
          
         
       
     
     The method further comprises transitioning to generally horizontal flight, as is shown in the second through fourth depictions in  FIG. 14-A , by adjusting the pitch of the blades  528  of rotor assemblies  511 ,  512  to generate forces perpendicular to longitudinal axis  508  until the rotorcraft attains an orientation in which longitudinal axis  508  is generally horizontal. Generally, when transitioning to horizontal flight it is desirable, but not necessary, to do so in a manner that results in the point  515  of the top side  514  of fuselage  502  being opposite the ground and the bottom side  516  of fuselage  502  generally facing the ground. Transition is preferably accomplished by transitioning blade angles from angles appropriate for vertical hover to angles appropriate for horizontal flight, examples of which are shown herein and made clear to those of ordinary skill in the art by the explanations and figures included herein. For illustration purposes, typical blade angles of the blades  528  of forward rotor assembly  511  and aft rotor assembly  512  at the mid point of the transition from vertical to horizontal flight (second depiction in  FIG. 14-A ) are summarized in the following table. 
     
       
         
               
             
               
               
             
               
               
               
               
               
               
               
               
               
             
               
               
               
               
               
               
               
               
               
             
           
               
                   
               
               
                 Vertical Orientation: Moving forward, craft pitch = 25° 
               
             
          
           
               
                   
                 Rotation angle of blade 
               
             
          
           
               
                   
                 0 
                 45 
                 90 
                 135 
                 180 
                 225 
                 270 
                 315 
               
               
                   
                   
               
             
          
           
               
                 FWD rotor blade angle 
                 35 
                 41 
                 43 
                 41 
                 35 
                 29 
                 27 
                 29 
               
               
                 AFT rotor blade angle 
                 35 
                 41 
                 43 
                 41 
                 35 
                 29 
                 27 
                 29 
               
               
                   
               
             
          
         
       
     
     The method further comprises the step of flying in a generally horizontal orientation by adjusting the pitch of blades  528  of rotor assemblies  511 ,  512  such that each blade  528  has a greater blade angle when rotating from fuselage top side  514  to fuselage bottom side  516  than when blade  528  is rotating from fuselage bottom side  516  to fuselage top side  514 . While in horizontal flight, the tips of rotor blades  528  describe long helical spirals through the air. Accordingly, the airspeed seen by rotor blades  528 , when the rotational speed of blades  528  is such that the blade tips are tracing arcs at 25 degrees from the horizontal, the blade tip speeds are only 10% faster than the speed of rotorcraft  501  itself. Rotorcraft  501  typically obtains lift from both down-going and up-going rotor blades  528 , with proportionally more lift being generated by the down-going blades  528 . As is well understood by those of ordinary skill in the art, the lift from a given rotor blade  528  is increased by increasing the angle of attack of the rotor blade  528  relative to the velocity of the approaching air. For illustration purposes only, typical blade angles of forward rotor assembly  511  and aft rotor assembly  512  during horizontal flight are summarized in the following table. 
     
       
         
               
             
               
               
             
               
               
               
               
               
               
               
               
               
             
               
               
               
               
               
               
               
               
               
             
           
               
                   
               
               
                 Horizontal Orientation: Moving forward, craft pitch = 90° 
               
             
          
           
               
                   
                 Rotation angle of blade 
               
             
          
           
               
                   
                 0 
                 45 
                 90 
                 135 
                 180 
                 225 
                 270 
                 315 
               
               
                   
                   
               
             
          
           
               
                 FWD rotor blade angle 
                 70 
                 72 
                 73 
                 72 
                 70 
                 68 
                 67 
                 68 
               
               
                 AFT rotor blade angle 
                 70 
                 72 
                 73 
                 72 
                 70 
                 68 
                 67 
                 68 
               
               
                   
               
             
          
         
       
     
     The method further comprises the step of transitioning back to generally vertical flight by adjusting the pitch of blades  528  to generate forces perpendicular to longitudinal axis  508  until rotorcraft  501  attains a generally vertical orientation, as is shown in  FIG. 14-B . Transition is preferably accomplished by transitioning blade angles from angles appropriate for horizontal flight back to angles appropriate for vertical hover. This is essentially the reverse of transition from vertical hover to horizontal flight described in greater detail above with respect to  FIG. 14-A . 
     The method further comprises the step of landing in a generally vertical orientation. Landing is essentially the reverse of the take off step described more fully above with respect to  FIG. 14-A . 
     The step of transitioning to horizontal flight in the method may further comprise additional steps. The additional steps comprise adjusting the pitch of blades  528  to generate horizontal velocity in the desired direction of travel as is illustrated in  FIG. 18 , further adjusting the pitch of blades  528  to cause forward fuselage end  504  to become oriented toward the desired direction of travel as is illustrated in  FIG. 17 , and further increasing velocity until rotorcraft  501  attains a generally horizontal orientation as is illustrated in  FIG. 16 . For suitably sized vehicles, the transition may typically be accomplished in this manner when the blade speed relative to the wind at the 0.7 radius of blade  528  is 313 feet per second and the lift coefficient is 0.8. Similarly, and as is illustrated in  FIG. 18 , horizontal velocity while in vertical hover may be attained without substantially changing orientation by adjusting the pitch of rotor blades  528  such that the blades  528  of forward rotor assembly  511  and the blades  528  of aft rotor assembly  512  generate proportionally more lift when they are diametrically opposed and moving away from and oriented perpendicular to the desired direction of travel. For illustration purposes only, typical blade angles of the blades of forward rotor  511  and the blades of aft rotor  512  needed to generate horizontal forces in the direction of top fuselage side  515  during vertical hover are summarized in the following table. 
     
       
         
               
             
               
               
             
               
               
               
               
               
               
               
               
               
             
               
               
               
               
               
               
               
               
               
             
           
               
                   
               
               
                 Vertical Orientation: Moving toward top side of fuselage, craft 
               
               
                 pitch = 0° 
               
             
          
           
               
                   
                 Rotation angle of blade 
               
             
          
           
               
                   
                 0 
                 45 
                 90 
                 135 
                 180 
                 225 
                 270 
                 315 
               
               
                   
                   
               
             
          
           
               
                 FWD rotor blade angle 
                 15 
                 18 
                 19 
                 18 
                 15 
                 12 
                 11 
                 12 
               
               
                 AFT rotor blade angle 
                 15 
                 18 
                 19 
                 18 
                 15 
                 12 
                 11 
                 12 
               
               
                   
               
             
          
         
       
     
     Those of ordinary skill in the art will recognize that additional steps may be added to the above embodiments of the flight method, including without limitation transitioning between various flight orientations and directions, and between stationary hovering and forward velocity, and that variations including such additional steps are also effective and are also within the scope of the present invention. 
     The rotorcraft  1  of the present invention exhibits several desirable flight characteristics. When hovering in a vertical orientation, the rotorcraft has a significant maneuvering advantage over traditional vehicles such as helicopters. In a helicopter, horizontal forces are realized by tilting the rotor axis. As is illustrated in relation to  FIG. 18 , rotorcraft  1  can develop horizontal forces without tilting its axis by adjusting continuously variable pitch blade assemblies  10  such that blades  28  of forward rotor assembly  110  produce additional lift when oriented 90 degrees from the desired direction of travel, and blades  28  of aft rotor assembly  210  produce a substantially equal amount of additional lift when also oriented 90 degrees from the desired direction of travel, thereby producing side force without changing vehicle attitude. Similarly, rotorcraft  1  can rotate about its longitudinal axis  8  by varying the collective pitch between rotor assemblies  110  and  210  so that their torques do not cancel. The result is excellent maneuverability when hovering, the ability to quickly adjust to wind gusts and turbulence, reasonable total power requirements, reasonable transition power requirements, and a small total rotor blade surface area. 
     The rotorcraft of the present invention has several practical applications. As an unmanned vehicle, it is well suited to applications including, but not limited to, automated border patrol and surveillance, identification of “hot spots” in dense forest fires, automated package delivery, automated aerial photography, and sky writing. In larger embodiments, the rotorcraft of the present invention may be adapted to carry passengers and cargo quickly and efficiently. 
     While the foregoing describes preferred embodiments of the rotorcraft and flight method of the present invention, it is to be understood that this description is to be considered as illustrative of the principles of the invention and is not to be limitative thereof, as numerous other variations, all within the scope of the invention, will readily occur to others of ordinary skill in the art.