Patent Publication Number: US-11649051-B2

Title: Rotary wing vehicle

Description:
RELATED APPLICATIONS 
     This application is a continuation of U.S. application Ser. No. 15/995,419, filed Jun. 1, 2018, which is a continuation of U.S. application Ser. No. 15/175,161, filed Jun. 7, 2016, which is a continuation of U.S. application Ser. No. 13/899,252, filed May 21, 2013, which is a continuation-in-part of U.S. application Ser. No. 13/270,872, filed Oct. 11, 2011 (now U.S. Pat. No. 8,469,307, issued Jun. 25, 2013) which is a divisional of U.S. application Ser. No. 12/872,622, filed Aug. 31, 2010 (now U.S. Pat. No. 8,042,763, issued Oct. 25, 2011), which is a divisional of U.S. application Ser. No. 11/105,746, filed Apr. 14, 2005 (now U.S. Pat. No. 7,789,341, issued Sep. 7, 2010), each of which are incorporated by reference herein in their entirety. This application also claims priority under 35 U.S.C. § 119(e) to U.S. Provisional Application Ser. No. 61/649,741, filed May 21, 2012 and U.S. Provisional Application Ser. No. 61/799,878, filed Mar. 15, 2013, both of which are incorporated by reference herein in their entirety. This application is also related to U.S. Provisional Application No. 60/562,081, filed Apr. 14, 2004, which is incorporated by reference herein in its entirety. 
    
    
     BACKGROUND 
     The present disclosure relates to aerial vehicles and particularly to unmanned aerial vehicles (UAV). More particularly, the present disclosure relates to unmanned rotary wing vehicles. 
     Rotary wing vehicles are used in a variety of applications. Unmanned rotary wing vehicles are often used by the military, law enforcement agencies, and commercial activities for aerial reconnaissance operations. 
     SUMMARY 
     A rotary wing vehicle, in accordance with the present disclosure includes a body structure having an elongated tubular backbone or core and a counter-rotating coaxial rotor system having rotors with each rotor having a separate motor to drive the rotors about a common rotor axis of rotation. A power source comprising, for example, a battery, fuel cell, or hybrid gas-electric generator is provided to supply electric power to the motors. Power transmission to and between the rotor systems is accomplished primarily by means of electrical wiring instead of mechanical shafting. A modular structure is described which assists manufacturability. 
     In illustrative embodiments, a torque tube is provided to transmit mechanical power inside the non-rotating tubular backbone creating a modular mast structure that can be used to support coaxial rotor systems on many types of vehicles. 
     In illustrative embodiments, a blade pitch control system is located between the rotor blades. A fixed, non-rotating body shell or aerodynamic fairing may be provided between the upper and lower rotors to protect the pitch control system and airframe against the elements and to reduce aerodynamic drag of the aircraft. 
     In illustrative embodiments, an auxiliary power-pack is provided which is separable from the vehicle in flight to facilitate, for instance, delivery of the vehicle to a distant location. In another embodiment, the power-pack comprises a payload such as an explosive munition, dipping sonar, hydrophones, or a separable sonobouy module. While aspects of the disclosure are applicable to many helicopters, including full-sized man carrying helicopters, the current disclosure is especially well suited for application to autonomous or radio-controlled rotary wing aircraft known as remotely piloted vehicles (RPVs), or unmanned aerial vehicles (UAVs). 
     Additional features of the present disclosure will become apparent to those skilled in the art upon consideration of illustrative embodiments exemplifying the best mode of carrying out the disclosure as presently perceived. 
    
    
     
       BRIEF DESCRIPTIONS OF THE DRAWINGS 
       The detailed description particularly refers to the accompanying figures in which: 
         FIG.  1    is a diagrammatic view of a rotary wing vehicle in accordance with the present disclosure showing an aircraft including a guidance system, and a pair of rotor systems coupled to an airframe comprising a non-rotating structural spine or backbone and carrying a payload; 
         FIG.  2 A  is a perspective view of a rotary wing vehicle in accordance with the present disclosure showing a counter-rotating coaxial rotor system in a vertical flight mode; 
         FIG.  2 B  is a perspective view of the rotary wing vehicle of  FIG.  2 A  having a counter-rotating coaxial rotor system and a fixed-wing booster module in a horizontal flight mode; 
         FIG.  3    is a side elevation view of the rotary wing vehicle of  FIG.  2 A  showing exterior body panels, electrical wiring, and booster section removed for clarity; 
         FIG.  4    is a side elevation view, with portions broken away, of the vehicle of  FIG.  2 A  showing a counter-rotating coaxial rotor system and an electrical power source; 
         FIG.  5    is an enlarged perspective view of the vehicle of  FIG.  2 A , with portions broken away, showing an upper interior section of the vehicle and the counter-rotating coaxial rotor system; 
         FIG.  6    is an enlarged perspective view of the vehicle of  FIG.  2 A , with portions broken away, showing a lower interior section of the vehicle and the counter-rotating coaxial rotor system; 
         FIG.  7 A  is a perspective view of a core tube or backbone having a circular cross section and a hollow interior channel that is used as a conduit between sections of the vehicle and showing electrical wiring running through the hollow interior and entering and exiting at various points; 
         FIG.  7 B  is a perspective view of backbone having a generally cruciform cross section with exterior channels running the length of the backbone that can be used as conduits between sections of the vehicle. 
         FIG.  8    is an enlarged perspective view of a first ring mount; 
         FIG.  9    is an exploded perspective view of a second ring mount showing attached linkages and body supports; 
         FIG.  10    is an enlarged perspective view of a middle interior section of the vehicle of  FIG.  2 A , with portions broken away, showing the counter-rotating coaxial rotor system; 
         FIG.  11 A  is an exploded perspective view of a rotor module having rotor blades with variable cyclic pitch and fixed collective pitch; 
         FIG.  11 B  is an exploded perspective view of a rotor module having rotor blades with variable cyclic and variable collective pitch; 
         FIGS.  12 A and  12 B  are perspective views of a first side and a second side of a motor mount; 
         FIGS.  13 A and  13 B  are perspective views of a first side and a second side of a rotor hub; 
         FIG.  14    is a sectional view taken along lines  14 - 14  of  FIG.  2 B , showing the rotor module; 
         FIG.  15    is a side elevation view of the counter-rotating coaxial rotor system of  FIG.  2 A , and a core tube depending from the rotor system; 
         FIGS.  16 A and  16 B  are exploded perspective views of a single power module including several batteries; 
         FIG.  17    is an orthographic view of the booster module of  FIG.  2 B  showing one wing folded for storage and one wing extended in a flight configuration; 
         FIG.  18 A  is an orthographic view depicting a rotatory wing vehicle in flight after separation from the booster module; 
         FIG.  18 B  is an orthographic view depicting the booster module after separation from the rotary wing vehicle of  FIG.  18 A ; 
         FIG.  19    is an elevation view of the rotary wing vehicle showing a dipping sonar or hydrophone assembly depending from a bottom portion of the vehicle; 
         FIGS.  20 A,  20 B, and  20 C  are sequential views of the rotary wing vehicle showing the operation of unequal length folding blades during a crash landing of the vehicle on ground underlying the rotary wing vehicle; 
         FIGS.  21 A and  21 B  are side elevation views of a storage tube and the rotary wing vehicle showing the vehicle folded for storage; 
         FIG.  22    is a perspective view of a rotary wing vehicle in accordance with present disclosure delivering a sensor or marking to a remote location shown for the purpose of illustration to be a ship on the open ocean; 
         FIG.  23    is a side elevation view of a rotary wing vehicle folded for storage in a rear portion of a gravity-delivered bomb; 
         FIG.  24    is a perspective view of a rotary wing vehicle deploying from the rear of a gravity-delivered bomb to the vicinity of a target site showing the gravity-delivered bomb ejecting the rotary wing vehicle and the rotary wing vehicle deploying into a vertical flight mode to loiter in the target area to provide an attacking force with real-time battle damage assessment after the gravity delivered bomb has struck the target; 
         FIG.  25 A  is a diagrammatic view of another rotary wing vehicle showing an aircraft having a central buss architecture with power and signal conduits, a guidance system, and a pair of rotor systems coupled to an airframe comprising a non-rotating structural spine or backbone and carrying a payload; 
         FIG.  25 B  is a diagrammatic view of the rotary wing vehicle of  FIG.  25 A  showing a rotor system, control system, and power supply communicating through a central data/power buss with power and signal conduit; 
         FIG.  26    is a diagrammatic view of another embodiment of a rotary wing vehicle, according to the present disclosure, having a central buss architecture with power and signal conduits, a guidance system, and a pair of rotor systems coupled to an air frame; 
         FIG.  27    is an elevation view of a rotary wing vehicle according to the present disclosure showing the rotary wing vehicle includes a streamlined body suited to high-speed translational flight and a coaxial mast module that includes an internal torque tube for driving an upper rotor; 
         FIG.  28    is an elevation view of the rotary wing vehicle of  FIG.  27    with portions of the body shells broken away to reveal the mast module and rotor control systems; 
         FIG.  29 A  is an enlarged side elevation view of the rotary wing vehicle of  FIG.  28    with portions of the mast module and rotor shroud cut away to reveal interior detail; 
         FIG.  29 B  is an enlarged portion taken from the circled region of  FIG.  29 A ; 
         FIG.  30    is an elevation view of another embodiment of a rotary wing vehicle in accordance with the present disclosure showing the rotary wing vehicle includes a streamlined body suited to high-speed translational flight and a coaxial mast module that includes an upper rotor speed reducer and showing that portions of body shells included in the streamlined body have been broken away to reveal a mast module and rotor control systems; 
         FIG.  31    is an enlarged elevation view of the rotary wing vehicle of  FIG.  29    with portions of the mast module and rotor shroud broken away to reveal interior detail; 
         FIG.  32    is a sectional view of the mast tube of the rotary wing aircraft of  FIG.  28   ; 
         FIG.  33    is an enlarged perspective view of a servo module included in a rotary wing vehicle showing that the servo module includes three servo actuators and three Z-links for varying the pitch of the upper and lower rotors at different phase angles simultaneously; 
         FIG.  34    is an enlarged perspective view of two pitch controller swashplates included in the servo module of  FIG.  33    showing the pitch controller swashplates connected by a Z-link to actuate the swashplates at different phase angles; 
         FIG.  35    is a plan view the swashplates and Z-link of  FIG.  34    showing a swashplate phase angle of about 90 degrees; 
         FIG.  36    is an exploded assembly view of the Z-link pitch control linkage of  FIGS.  33  and  34   ; 
         FIG.  37    is plan view of a rotary wing vehicle in accordance with the present disclosure showing an upper rotor phase angle (solid double arrow) and a lower rotor phase angle (hollow double arrow) and a resulting total rotor system phase angle (combined solid and hollow double arrow); 
         FIG.  38    is a side elevation view of a rotorcraft power and control system according to the current disclosure configured for an aircraft with a single drive motor, two rotors and a pusher propeller; 
         FIG.  39    is an enlarged perspective view of the rotorcraft of  FIG.  38    showing details of the main shaft splitter and drive gears for the counter-rotating rotors and the belt-drive system for the pusher propeller; 
         FIG.  40    is a perspective end view of a main rotor mast configured with internal passageways for a torque tube and electrical wiring or plumbing; 
         FIG.  41    is a perspective end view of a main rotor mast configured with internal passageways for a torque tube and six mechanical slider linkages; 
         FIG.  42    is a perspective view of a main rotor mast assembly including the main rotor mast of  FIG.  41    and six slider linkages engaging the six interior passageways and connected to upper and lower swashplates; 
         FIG.  43    is an enlarged perspective end view of the main rotor mast assembly of  FIG.  42    showing six swashplate slider linkages engaging the six interior mast passageways; 
         FIG.  44 A  is a perspective view of a slider linkage configured with a downward pointing follower link to control a lower swashplate; 
         FIG.  44 B  is an exploded perspective view of a slider linkage configured with an upward pointing follower link to control an upper swashplate; 
         FIG.  45    is a perspective side view of a helicopter with a non-rotating mast and six rotary servo actuators coupled to the mast with upper and lower rotor hubs and rotor blades removed for clarity; 
         FIG.  46    is an enlarged perspective end view of the non-rotating main rotor mast assembly of  FIG.  45    showing the six rotary servo actuators coupled to the mast and connected to the upper and lower swashplates with six individual linkages; 
         FIG.  47    is a perspective view of a high-speed helicopter in accordance with the present disclosure showing that the high-speed helicopter includes a non-rotating mast supporting an aerodynamic mask shroud between the upper and lower rotor blades to reduce drag; 
         FIG.  48    is an enlarged partial perspective side view of the helicopter of  FIG.  47    with portions broken away to reveal the non-rotating mast, mast shroud, six linear servo actuators, and other control system components including electronics and antennae supported by the mast between the upper and lower rotor blades; 
         FIG.  49    is an enlarged partial perspective view of the non-rotating mast assembly of the helicopter  FIG.  47    showing upper and lower rotor hubs, upper and lower rotor drive gears, and linear servo actuators; 
         FIG.  50    is a perspective view of the central non-rotating mast of the mast assembly shown in  FIG.  49    with the mast sleeve removed to show details of the electrical bus inlays; 
         FIG.  51    is a sectional view taken along line A-A of  FIG.  50    showing a torque tube inside the mast and showing exterior channels for electrical bus inlays; 
         FIG.  52    is a perspective view of the electrical bus inlays of  FIG.  51   ; 
         FIG.  53    is a perspective view of a mast sleeve with six interleaved linear servo actuators and two swashplates configured to reduce a mast assembly frontal area; and 
         FIG.  54    is a plan view of the lower swashplate of  FIG.  53    showing the relationship between the swashplate arms to reduce the frontal area of the mast assembly. 
     
    
    
     DETAILED DESCRIPTION 
     As suggested diagrammatically in  FIG.  1   , a rotary wing vehicle  1  includes, in series, a first module  2 , a first and a second rotor system  3 ,  5 , power modules  13  and  14 , and a second module  15  coupled in spaced-apart relation to an airframe  40  extending along a common axis  7 . Illustratively, airframe  40  is an elongated central backbone  40  and can be arranged as a hollow core or having a cruciform cross-section. In operation, first rotor system  3 , also called first rotor  3 , and second rotor system  5 , also called second rotor  5 , rotate in opposite directions about common axis  7  to direct thrust in direction  24  and create lift in direction  24 ′ to cause controlled flight of rotary wing vehicle  1 , as suggested in  FIG.  2 A . First module  2  is adapted to include a variety of guidance systems  50 ′, electronics  55 , or payloads  15 ′. Second module  15  is adapted to include payload  15 ′, or in some embodiments, a variety of guidance systems  50 ′ and electronics systems  55 ′. Payload  15 ′ may include, but is not limited to, munitions, radiation sensors, chemical detection sensors, biological agent sensors, active and passive listening devices, video sensors, supplemental power sources, or other mission-specific equipment. Rotary wing vehicle  1  thus provides means for moving reconnaissance, observation, or survey monitoring equipment to an area of interest to obtain information therefrom. 
     As suggested in  FIGS.  1 ,  25 A, and  25 B , first rotor system  3  includes a first motor  54 , first rotor blades  20 , and a first pitch controller  56 . In illustrative embodiments, motor  54  is an electric motor as shown, for example, in  FIGS.  4 - 6   , or other suitable means for providing power to rotate rotor blades  20  about common axis  7 . First rotor system  3  and second rotor system  5  are similar to one another in structure and function. Second rotor system  5  includes a second motor  61 , second rotor blades  22 , and a second pitch controller  57 . In illustrative embodiments, motor  61  is an electric motor as shown, for example, in  FIGS.  4 - 6   , or other suitable means for providing power to rotate rotor blades  22  about common axis  7 . Illustratively, electrical and electronic components are connected and communicate through electrical conduit  173  and electronic conduit  174  which hold power and signal lines, respectively. Although rotary wing vehicle  1  is illustrated having two rotor systems, rotary wing vehicle  1  may have more than two rotor systems as performance and mission demands dictate. 
     As shown in  FIGS.  1  and  3   , airframe  40  is non-rotating and forms a central elongated hollow backbone to receive first module  2 , first and second rotor systems  3 ,  5 , power modules  13  and  14 , and second module  15 . Illustratively, power modules  13  and  14  are positioned to lie in side-by-side relation to one another between second rotor system  5  and second module  15 . Because airframe  40  is hollow power modules  13 ,  14  can be connected electrically through the hollow backbone to motors  54  and  61 . 
     Illustratively, pitch controller  56  is a swashplate  56 ′ coupled to a fore/aft servo  58  and a roll servo  59  to vary the cyclic pitch of rotor blades  20  in response to input from a controller  55 . In some embodiments, swashplate  56 ′ is further coupled to a collective servo  98  to collectively change the pitch of rotor blades  20 . Likewise, pitch controller  57  is a swashplate  57 ′ coupled to a fore/aft servo  58  and a roll servo  59  to vary the cyclic pitch of rotor blades  20  in response to input from a controller  55 . In some embodiments, swashplate  57 ′ is also coupled to a collective servo  98  to collectively vary the pitch of rotor blades  20 . In illustrative embodiments, controller  55  is a command signal controller as shown, for example, in  FIG.  3   , or other suitable means for providing a desired electrical or mechanical directional signal to servos  58 ,  59 , or  98 , and motors  54 ,  61 . 
     Illustratively, rotary wing vehicle  1  has a fixed-pitch rotor system having two servos  58 ,  59  for aircraft pitch (helicopter-style fore/aft cyclic input) or aircraft roll (helicopter-style right/left cyclic input) control. Servo  98 , shown in phantom in  FIG.  1   , can be mounted similarly to servos  58 ,  59  if collective pitch control is desired. In embodiments having a fixed-pitch rotor system, rotor systems  3 , 5  are connected to swashplates  56 ′,  57 ′ by pitch links  119 . Servos  58 ,  59  are connected to swashplates  56 ′,  57 ′ by pitch links  125 ,  126 . A feature of the present disclosure is that rotary wing vehicle  1  can be flown with as few as one or two cyclic servo actuators (servo  58 ,  59 ). In a “one-servo” flight mode, differential torque of motors  54 ,  61  controls yaw orientation, and servo  58  controls forward and backward flight. With only one cyclic servo, rotary wing vehicle  1 , also called vehicle  1 , can be flown much like an airplane having only rudder and elevator control. In the illustrative “two-servo” flight mode, servos  58 ,  59  provide fore/aft aircraft pitch and right/left aircraft roll control with differential torque of motors  54 ,  61  providing yaw control. 
     In operation, rotor hubs  101  rotate in opposite directions. Servos  58 ,  59  are controlled by onboard flight control electronics to tilt simultaneously swashplate  56 ′ and swashplate  57 ′ which then cyclically vary the blade pitch angle of rotating rotor blades  20  to tilt vehicle  1  in one of aircraft pitch direction  170  and aircraft roll direction  171 . In another embodiment having collective pitch (see  FIG.  11 B ), collective servo  98  and a third pitch link (not shown) are provided to vary the axial location of swashplates  56 ′,  57 ′ along common axis  7  and to vary the collective pitch of rotor blades  20 ,  22  using electronic Collective-Cyclic Pitch Mixing (CCPM). With collective-cyclic pitch mixing servos  58 ,  59 , and  98  tilt swashplates  56 ′ and  57 ′ in unison to vary cyclic pitch and move swashplates  56 ′,  57 ′ axially in unison along common axis  7  to vary collective pitch. 
     The illustrative embodiment employs differential motor speed for yaw (heading) control while in a vertical flight configuration. Normally, coaxial helicopters use variable blade pitch and differential blade angle to control yaw motions in flight. In the present disclosure, differential torque generated by operating motors  54 ,  61  at different speeds relative to the fixed body of vehicle  1  generates yaw forces to stabilize and control yaw motion (i.e. rotation about common axis  7 ). In this method, the torque (and eventually the speed) of motor  54  is increased or decreased in response to a yaw motion of rotary wing vehicle  1  about vertical common axis  7 . The torque (speed) of second motor  61  is adjusted automatically by an onboard computer system, contained within controller  55 , in opposition to the torque (speed) of first motor  54  to maintain constant lift so that rotary wing vehicle  1  neither gains nor loses altitude. 
     Rotor blades  20  and  22  are coupled to rotary wing vehicle  1 , also called rotary wing aircraft  1 , and supported for rotation by rotor hubs  101 . Rotor hubs  101  are further coupled for pivotable movement to an internal yolk  108 , as shown best in  FIG.  11 A . Pivot axles  109  extend through rotor hub  101  and are received by yolk  108 . Yolk  108  is adapted to couple a pair of rotor blades to rotor hub  101  for rotation about common axis  7 . Yolk  108  is further coupled to a first end of a pair of pitch links  119 . Each pitch link  119  is further coupled on a second end to a perimeter edge of swashplate  56 ′ or  57 ′. Thus, yolk  118  is pivoted by input from swashplate  56 ′,  57 ′ in response to linear motion input from servos  58 ,  59 , or  98 . This pivoting motion of yolk  118  in turn causes each rotor blade  20 ,  22  to pivot in response, thus increasing or decreasing the rotor blade pitch of rotor blades  20 ,  22 . 
     As suggested in  FIGS.  2 A and  2 B , a rotary wing vehicle  1  includes an upper section  2 ′, first and second rotors  3  and  5 , a middle section  4 , a lower section  6 , first and second power modules  13 ,  14 , and a payload  15 ′ arranged in spaced apart relation along common axis  7 . Referring now to  FIGS.  2 A- 4   , internal mechanical and electrical components within upper section  2 ′ and middle section  4  of vehicle  1  are enclosed by a thin-walled upper body shell  10  and a middle body shell  11 , respectively. A lower body shell  12  covers a portion of lower section  6 , but could be extended to cover all of lower section  6 . A feature of the present disclosure is that body shells  10 ,  11  are blow-molded from a plastic material such as polycarbonate or ABS, and, in conjunction with backbone  40 , form a structure for rotary wing aircraft that has both a central strength component and a thin exterior cover component that together are stiff, strong and easy to manufacture. 
     As shown in  FIG.  3   , a rotary wing aircraft  1  in accordance with the present disclosure has a rotor system comprising a motor  54  operably connected to rotor blades  20  by means of a drive train such as gears  106 ,  107  ( FIG.  11   ). A pitch control such as a swashplate  56 ′ ( FIG.  10   ) is operably connected to rotor blades  20  to vary the cyclic and/or collective pitch of rotor blades  20  in response to output from a servo actuator such as servos  58 , 59  ( FIG.  3   ) through linkages such as pitch links  125 ,  126  ( FIG.  10   ). Power such as electricity from batteries (not shown) or fuel from a storage tank (not shown) in a power module  13  flows through a power conduit across rotor system and provides power to operate controller  55 , motor  54 , and servos  58  and  59 . Control signals from controller  55  flow along a signal conduit and regulate the speed of motor  54  and the positioning output of servos  58  and  59 . The power conduit and signal conduit are conducted between an inflow side and an outflow side of rotor blades  20  through channels  96 , also called interior space  96 , formed in the structural spine or backbone  40  ( FIGS.  7 A,  7 B, and  15   ) of vehicle  1 . 
     In hovering flight, first rotor  3  and second rotor  5  rotate in opposite directions about common axis  7  forcing air downward in direction  24  and lifting vehicle  1  in an upwardly direction, as suggested in  FIG.  2 A . First rotor  3  has rotor blades  20  configured to rotate in direction  21 , and second rotor  5  has rotor blades  22  configured to rotate in direction  23  about common axis  7 . Because first rotor blades  20  and second rotor blades  22  are equipped with a cyclic pitch control, vehicle  1  is configured for directional flight in direction  25  wherein common axis  7  is orientated substantially vertically. 
     Referring now to  FIG.  2 B , a second embodiment contemplated by the current disclosure is depicted having a booster module  8  appended to lower section  6  at a booster interface  9 . Booster module  8  contains, for example, an auxiliary power source (not shown) to augment an internal power source contained in power modules  13  and  14  carried in vehicle  1 . Illustratively, the auxiliary power source (not shown) and power modules  13  and  14  are electrical batteries  13  and  14 . Booster module  8  includes left and right wings  16 ,  17  to provide additional lift for vehicle  1  in directional flight in direction  18  wherein common axis  7  is oriented substantially horizontally. 
     Airframe  40  forms a structural backbone of rotary wing vehicle  1  and generally runs vertically through the center of rotary wing vehicle  1  from upper section  2 ′ to lower section  6 , as shown best in  FIG.  4   . Illustratively, airframe  40  is a non-rotating core tube with a hollow interior channel  96  ( FIG.  7 A ) or a cruciform beam  97  with exterior channels ( FIG.  7 B ). First and second rotor systems  3 ,  5 , also called first and second rotor modules  3 ,  5 , all components within upper section  2 ′, middle section  4 , and lower section  6  are coupled to airframe  40 . Referring now to  FIG.  7 A , elongated central backbone  40 , also called non-rotating hollow core tube  40 , further acts as a conduit for electrical wiring  45 , plumbing (not shown), and mechanical linkages (not shown) passing between components in upper section  2 ′, middle section  4 , and lower section  6  of rotary wing vehicle  1 . Longitudinal slots  46  and  47  are provided as entry and exits points for electrical wires  45 , plumbing, and linkages. Since non-rotating hollow core tube  40  and cruciform beam are unitary and continuous between body sections  2 ,  4 , and  6 , the rigidity and light-weight structural properties of vehicle  1  are increased. Illustratively, non-rotating hollow core tube  40  and cruciform beam  97  are preferably made of wound or pultruded carbon graphite fiber, fiberglass, or aluminum alloy number  7075  (or similar) with an outside diameter (core tube  40 ) or width dimension (cruciform beam) of about 0.5 inches (13 mm) and a wall thickness of between about 0.03 inches (0.76 mm) and about 0.05 inches (1.3 mm). 
     Rotary wing vehicle  1  is arranged having three body sections, as shown best in  FIG.  3   . Upper section  2 ′ is arranged having a horizon sensor/stabilizer  50 , an electronic gyro stabilizer  51 , a gyro mounting table  52  coupled to an upper end of core tube  40 , a first motor speed controller  53 , a first motor  54 , a radio receiver, and controller  55 . Middle section  4  includes a first swashplate  56 ′, a second swashplate  57 ′, a fore-aft cyclic servo  58 , and a roll cyclic servo  59 . Lower section  6  includes a second motor speed controller  60 , a second motor  61 , a radio battery  62 , first and second power modules  13  and  14 , and payload module  15 . 
     In the illustrated embodiment, horizon sensor/stabilizer  50  is a model “FS8 Copilot” model by FMA company, electronic gyro stabilizer  51  is a “G500” model silicone ring gyro by JR company, motors  54 ,  61  are “B2041S” models by Hacker company, and motor speed controllers  53 ,  60  are “Pegasus  35 ” models by Castle Creations company which are computer-based digital programmable speed controllers. Rotary wing vehicle  1  is also configured to receive a GPS receiver/controller and telemetry system (not shown), arranged to be coupled to upper section  2 ′. 
     Interior components of rotary wing vehicle  1  are coupled to core tube  40  by ring mounts  70 , as shown in  FIG.  8   . Ring mount  70  includes an annular inner portion  71  conforming to the annular exterior surface of core tube  40 . Ring mount  70  includes radially extending mounting arms  72 ,  73 ,  74  having flanges  75 ,  76 ,  77  adapted to hold mechanical, electrical, and other interior components of rotary wing vehicle  1 . Ring mount  70  is arranged to support motor  54  in flange  75 , motor speed controller  53  on flange  76 , and radio receiver  55 ″ on flange  77 . Interior components of vehicle  1  are coupled, for example, to mounting flanges using a variety of fasteners (such a nylon ties through apertures  78 ) or adhesives. Annular portion  71  provides means for locking ring mount  70  to non-rotating hollow core tube  40  to prevent ring mount  70  from rotating or sliding axially along non-rotating hollow core tube  40 . Means for locking ring mount  70  to non-rotating hollow core tube  40  includes fasteners (not shown) received by set screw receiver  79  or a variety of adhesives. A second ring mount  80 , as shown in  FIG.  9   , includes an annular ring  63 , arms  82  and  83 , and axial posts  84 ,  85  for supporting body standoffs  86 ,  87 ,  88 , swashplate anti-rotation arms  90  and  91 , and swashplate links  92  and  93 . 
     Servo module  81  includes ring mount  80  supporting pitch servo  58 , roll servo  59 , and universal body standoffs  86 ,  87  (as described in U.S. Provisional Patent Application No. 60/525,585 to Arlton which is hereby incorporated by reference herein) which support middle body shell  11 , as shown, for example, in  FIG.  10   . As suggested in  FIGS.  3 ,  4 ,  5 ,  6 ,  9 ,  10  and  15   , body standoffs  86 ,  87 ,  88  are secured to ring mount  80 . Through-holes  263  in body standoffs  86 ,  87 ,  88  are receptive to many types of commercial fasteners such as bolts and rods (not shown) for securing body standoffs  86 ,  87 ,  88  to ring mount  80  and middle body shell  11 . Middle body shell  11  is generally secured to body standoffs  86 ,  87 ,  88  to provide a cover and aerodynamic fairing for servos  58 ,  59  and swashplates  56 ′,  57 ′. Ring mounts  70 ,  80  are arranged to incorporate and support many structural features of rotary wing vehicle  1 . Ring mounts  70 ,  80  assist assembly of rotary wing vehicle  1  because ring mounts  70 ,  80  and associated interior components can be preassembled as subassemblies and then later assembled along with other modules to non-rotating hollow core tube  40  in a final manufacturing step. 
     Referring now to  FIGS.  11 A,  12 A,  12 B,  13 A,  13 B and  14   , rotor system  3 , also called rotor module  3 , includes a rotor mount  100 , a rotor hub  101  having an internal gear  107 , first and second ball bearings  102  and  103 , a shaft  101 A extending between bearings  102  and  103 , a ring clip  104 , motor  54 , a planetary gearbox  105 , a pinion gear  106 , a blade yolk  108 , pivot axles  109 , axle end caps  110 , torsion springs  111 , and rotor blades  20 . A motor mount  122  is receptive to gearbox  105  to couple motor  54  to rotor mount  100 . When assembled, bearings  102 ,  103  are retained by ring clip  104  engaging slot  99  on a boss  112  extending from rotor mount  100 . Rotor blade  20  is held in place by a pin  113  extending through cap  110  and aperture  114  formed in axle  109 . Axle  109  passes through a bearing aperture  117  formed in rotor hub  101  and into an aperture  94  in yolk  108  when it is retained by another pin (not shown). Pitch links  119  couple yolk  108  to swashplate  56 ′. 
     As shown in  FIG.  11 B , a rotor module adapted to support both cyclically and collectively pitchable rotor blades includes collective rotor hub  201  that is similar to rotor hub  101  and receptive to a collective yolk frame  208  coupled to bosses  214  formed on an interior surface of hub  201  by fasteners  212 . Collective yolk frame  208  supports the radial flight loads produced by rotor blades  20  acting through thrust bearings  203 . Pitch links  119  couple pitch arms  210  to swashplate  56 ′. 
     Illustratively, planetary gearbox  105  has a reducing speed ratio of about 4:1. Pinion gear on motor  54  has nine teeth and engages internal gear  107  on rotor hub  101  which has sixty teeth, so the total speed reduction ratio of rotor module  3  is about 26.7:1 (that is, the output shaft of motor  54  turns 26.7 times for each turn of rotor hub  101 ). This reduction ratio encourages the use of high efficiency electric motors running at high voltages and high speeds. 
     Illustratively, motor  54  is a brushless motor. In some applications, especially where flight times are short and economy is a factor (for example, in a short-range disposable munition) several low-cost brushed motors (i.e. motors having carbon brushes and rotating commutators) are used in place of one high-cost brushless motor  54  to turn rotor hub  101 . In such cases, while rotor module  3  is shown having one motor  54  to drive rotor hub  101 , it is within the scope of this disclosure to include several motors around the circumference of rotor mount  100  to drive rotor hub  101  instead of only one. It is also anticipated that rotor hub  101  itself can be configured with wire coils and magnets to act as a motor so that no separate motors are required to drive rotor hub  101  about common axis  7 . 
     Rotor blade  20  in the embodiment shown is injection molded of polycarbonate plastic material and is of the type described in U.S. Pat. No. 5,879,131 by Arlton, which patent is hereby incorporated by reference herein. Rotor blade  20  is free to flap upward and downward about 6 degrees about flapping axis  120  before tabs  121  on torsion springs  111  contact pitch axle  109  and resist further flapping. This means that rotor blades  20  can flap up and down freely in flight about +/−6 degrees and can fold upward 90 degrees and downward 90 degrees for storage or during a crash landing. 
     In the embodiment shown in the drawings, rotor mount  100  is injection molded in one piece from a thermoplastic material such as polycarbonate or nylon. Rotor hub  101  is injection molded in one piece from a thermoplastic material such as nylon or acetal. Rotor blades  20  are supported in flight by rotor hub  101  (which forms part of the exterior body shell of vehicle  1  instead of by traditional coaxial shafts coincident with common axis  7 . This places rotor support bearings  102 ,  103  very close to rotor blades  20  and frees space within the central body portion of rotary wing vehicle  1  for other mechanical or electrical components. In a fixed-pitch rotor system (shown in the drawings) radial flight forces produced by rotating blades  20  are supported by internal yolk  108  which connects two rotor blades  20  and which includes an internal aperture surrounding and bypassing core tube  40 , thus no special thrust bearings are required. 
     Referring now to  FIG.  15   , a coaxial rotor system in accordance with the current disclosure comprises core tube  40 , two rotor systems  3 ,  5 , two swashplates  56 ′ and  57 ′, and one servo module  81  coupled to non-rotating hollow core tube  40  in mirrored symmetry around servo module  81 . While a coaxial rotor system with two rotors is disclosed, rotary wing vehicle  1  could be equipped with additional rotor systems (not shown) spaced apart along the length of non-rotating hollow core tube  40  for additional thrust or operational capabilities. 
     In the illustrated embodiment, rotary wing vehicle  1  has a fixed-pitch rotor system which requires only two servos  58 ,  59  for aircraft pitch (fore-aft cyclic) and aircraft roll (right-left cyclic) control. A third collective servo  98  can be mounted in a similar fashion in middle section  4 , for instance, if collective pitch control is desired. 
     Rotor systems  3 , 5  are connected to swashplates  56 ′,  57 ′ by pitch links  119 . Servos  58 ,  59  are connected to swashplates  56 ′,  57 ′ by pitch links  125 ,  126 . In operation, rotor hubs  101  rotate in opposite directions. Servos  58 ,  59  are controlled by onboard flight control electronics  55 ′ to tilt simultaneously swashplate  56 ′ and swashplate  57 ′ which then cyclically vary the blade pitch angle of rotating rotor blades  20  to tilt vehicle  1  in one of aircraft pitch direction and aircraft roll direction. In another embodiment having collective pitch (see  FIG.  11 B ), a third servo and third pitch link (not shown) are provided to vary the axial location of swashplates  56 ′,  57 ′ along common axis  7  and to vary the collective pitch of rotor blades  20 ,  22  using electronic Collective-Cyclic Pitch Mixing (CCPM). Using servos positioned to lie between rotor systems  3 ,  5  and directly coupling control swashplates  56 ′,  57 ′ with linkages to control a coaxial rotor system in this way is a feature of the embodiment. 
     An illustrative embodiment of the disclosure includes motors  54 ,  61  positioned to lie above and below rotor blades  20 ,  22  (see  FIG.  25 A ) with power transmission between the rotor systems  3 ,  5  accomplished through electrical wiring  45  instead of mechanical shafting thereby reducing mechanical complexity and weight. In another embodiment (see  FIG.  26   ), motors  54 ,  61  are positioned to lie between the rotor blades  20 ,  22 , and servo actuators  58 ,  59  are positioned to lie in spaced-apart relation to locate rotor blades  20 ,  22  therebetween (see  FIG.  26   ). Because power and control of the rotor systems  3 ,  5  is entirely electrical in nature, the entire control system of rotary wing vehicle  1  can be operated electrically by digital computers and solid-state electronics without mechanical linkages or hydraulic amplification. Locating the motors  54 ,  61 , as shown in  FIG.  25 A , eliminates the need for concentric rotating shafting between rotor blades  20 ,  22 , and positions servos  58 ,  59  to drive both swashplates  56 ′, (included in first pitch controller  56 )  57 ′ (included in second pitch controller  57 ) directly. 
     A feature of the present disclosure is that vehicle  1  can be flown with as few as one or two cyclic servo actuators (servo  58 ,  59 ). In a one-servo flight mode, differential torque of motors  54 ,  61  controls yaw orientation, and servo  58  controls forward and backward flight. With only one cyclic servo, vehicle  1  can be flown much like an airplane having only rudder and elevator control. In a two-servo flight mode, as illustrated in the drawings, servos  58 ,  59  provide fore/aft aircraft pitch and right/left aircraft roll control with differential torque of motors  54 ,  61  providing yaw control. 
     In another embodiment of the current disclosure, power to drive motors  54 ,  61  in flight is provided by high-capacity electric batteries  130  such as lithium-polymer or lithium-ion batteries, or fuel cells. Referring now to  FIGS.  16 A and  16 B , power module  13  has six rechargeable lithium ion batteries  130  arranged in a hexagonal pattern around non-rotating hollow core tube  40  and wired in series to produce about 21.6 volts of electrical potential. Battery ring mount  131  is formed to include center aperture (ring)  132  to accommodate non-rotating hollow core tube  40  and flange  133  to hold batteries  130 . Electrical wires  45  from power module  13  enter non-rotating hollow core tube  40  at opening  47  (see  FIG.  7 A ), and are routed through non-rotating hollow core tube  40  to motor speed controllers  53 ,  60 . 
     As shown best in  FIG.  25 A  multiple power modules  13 ,  14  are provided for additional energy capacity during flight and are, illustratively, wired in parallel to increase the electrical current available to motors  54 ,  61 . Flight times of rotary wing vehicle  1  can be adjusted by adjusting the number of power modules  13 ,  14  carried in flight. 
     Extra locking rings (or ring mounts with no radial arms)  135  are provided above and below power module  13 ,  14  to help couple power modules  13 ,  14  to non-rotating hollow core tube  40 , as shown, for example, in  FIG.  4   . Since power modules  13 ,  14  are relatively heavy compared to other components of vehicle  1 , locking rings  135  prevent power modules  13 ,  14  from sliding along non-rotating hollow core tube  40  during a crash landing of rotary wing vehicle  1 . A feature of the present disclosure is that rotary wing vehicle  1  is well-suited to be manufactured and assembled in modules. Rotor, wing, control, power, booster, electronics, and payload modules are manufactured separately and slid onto core tube  40 . Electrical connectors for connections passing through openings  46 ,  47  in core tube  40  are mounted flush with the surface of core tube  40  to assist in assembly and disassembly of vehicle  1  for maintenance and repairs. 
     Energy density and power density are considerations in UAV design and can be applied to an aircraft as a whole. Aircraft with higher energy densities and power densities have better overall performance than aircraft with lower densities. In general, energy density and power density are defined as the amount of energy and power available per unit weight. For example, the energy density of a fuel or electric battery (also known as “specific energy”) corresponds to the amount of energy contained in a unit measure of fuel or battery (measured, for instance, in Nm/Kg or ft-lbs/slug). 
     Chemical (liquid) fuels tend to have higher energy densities than electric batteries. One additional characteristic of liquid fuel power as compared to electric battery power is that the weight of a liquid fueled aircraft decreases over the course of a flight (as much as 60%) as it burns fuel. Consequently the energy density of a liquid fueled aircraft (i.e., the energy available per unit weight of the aircraft) decreases slowly and power density (power available per unit weight) increases as it flies. This means that the performance of liquid fueled aircraft actually improves near the end of a flight. 
     In contrast, the overall power density of an electric-powered aircraft is constant throughout the flight because the maximum output power of the batteries is almost constant and the batteries do not lose weight as they discharge. Energy density also decreases quickly because the total energy available decreases. To improve energy and power density of the current disclosure, an auxiliary electric booster or power module  8  is provided that can be jettisoned in flight after its energy supply is depleted. Thus, booster module  8  comprises additional battery modules (not shown) assembled around common axis  7  with a mechanism to retain booster module  8  to rotary wing vehicle  1 . 
     In another embodiment, booster module  8  includes an internal combustion engine (such as a diesel engine not shown) which drives an electric generator (not shown) to convert chemical energy contained in a chemical fuel to electrical energy. In other embodiments contemplated by this disclosure, a turbo-electric generator system (not shown) may be used to create electrical energy. A consideration of a booster module  8  containing such a gas-electric generator is that the entire weight of the module, fuel system, and engine, can be jettisoned at the end of a first flight phase leaving the relatively low weight rotary wing vehicle  1  to complete a second flight phase. 
     In the illustrative embodiment, booster module  8  includes foldable wings  16 ,  17  to increase lift in a horizontal flight mode of rotary wing vehicle  1 . As shown in  FIG.  17   , wing  17  is folded about folding axis  140  for compact storage. Wings  16 ,  17  are attached at about their “quarter chord” location to pivot shafts (not shown). When deployed for flight with pivot shafts held rigidly perpendicular to common axis  7  (see also  FIG.  2   ), wing  16  is free to pivot about pitch axis  143  to find its own best angle of attack. Because wings  16 ,  17  are free to rotate about their own pitch axes in flight, appendages such as wings  16 ,  17  are sometimes referred to as “free-wings.” It should be noted that wings  16 ,  17 , being free-wings, can operate efficiently over a wide speed range because of their ability to change pitch automatically to meet the oncoming airflow. Application of such a free wing to a rotary wing UAV is a feature of the disclosure. 
     In high-speed horizontal flight, common axis  7  is orientated substantially horizontally with rotor modules  3 ,  5  together acting like a single counter-rotating propeller to pull rotary wing vehicle  1  in a horizontal direction  18 . Wings  16 ,  17  help to lift lower section  6  and booster module  8  so that rotor modules  3  and  5  can apply more power to forward propulsion and less to vertical lifting. 
     It should also be noted that the current disclosure does not require aerodynamic control surfaces (such as on wings  16 ,  17 ) because cyclic control of rotor module  3 ,  5  provides control power for maneuvering in aircraft pitch (elevation) direction  144  and aircraft yaw (heading) direction  145  when common axis  7  is substantially horizontal. Airplane-style roll control (about common axis  7 ) during high-speed horizontal flight is accomplished though differential torque/speed of rotor modules  3 ,  5 . This method of control for horizontal flight of a rotary-wing UAV is a feature of the illustrative embodiment. 
     Referring now to  FIGS.  18 A and  18 B , when the energy of booster module  8  has been depleted, a command from on-board controller  55  of rotary wing vehicle  1  actuates a mechanism such as a latch (not shown) that separates booster module  8  from rotary wing vehicle  1  and booster module  8  falls away in direction  19 . Rotary wing vehicle  1  then, in one flight mode, assumes a more vertical orientation and flies like a helicopter. 
     In another embodiment, booster module  8  includes a mission-specific payload  147  such as an explosive munition, dipping sonar, hydrophones, radio ID marker, or a sonobouy. As illustrated in  FIG.  19   , upon separation from rotary wing vehicle  1 , booster module  8  falls away leaving a sonar or hydroponic system  147  or other sensor connected to rotary wing vehicle  1  by wire or fiber optic cable  146  so that rotary wing vehicle  1  can move payload  147  from place to place, deliver payload  147  accurately to a desired location, and act as a telemetry link between payload  147  and a remote receiver (not shown). This can be an effective method of, for example, monitoring a target or marking a ship at sea with a remote radio ID marker or other marking instrument. 
       FIG.  22    illustrates a method of delivering a marker comprising, for example, a sensor, or a marking device, such as indelible paint or a radio transmitter, to a remote location, in this case a ship on an open ocean  157 . Vehicle  1  is shown approaching ship S (in frame), maneuvering to touch ship S and leaving the marker on ship S (in frame) and exiting the area (in frame). This method of marking is a feature of the present disclosure that allows a point of interest to be monitored after vehicle  1  has left the local area. Alternatively or in conjunction, vehicle  1  can retain a sensor when it leaves the local area which may, for instance, have taken a sample of the atmosphere near ship S, and return the sensor and sample to a remote processing point for further analysis by a mass spectrometer, biological or radiological measuring device or other such device (not shown). While the point of interest shown in the drawings as a ship S, it will be understood that ship S could be any other point of interest accessible to vehicle  1  such as a truck, aircraft, building, tower, power line, or open area of land. 
     Another embodiment of the current disclosure shown in  FIGS.  20 A,  20 B , and  20 C, has unequal length folding, coaxial rotor blades  148 ,  149  with upper blades  148  having a greater span than lower blades  149 . This is a feature arranged so that during a crash landing of vehicle when upper blades  148  contact the ground  155  before lower, shorter blades  149  so that upper blades  148  fold away from, or faster than, lower blades  149  thereby reducing the possibility that upper blades  148  and lower blades  149  will contact each other while still rotating at high speed. As shown in the drawings, lower blades  149  span about 20 to 22 inches (51 cm to 56 cm). 
     The ability to fold for compact storage and for landing is another feature of the current disclosure. As shown in  FIGS.  21 A and  21 B , rotary wing vehicle  1  is compact enough to fit inside a standard A-size sonobouy tube used by the United States Navy. The unique core-tube structure of the current disclosure not only allows rotary wing vehicle  1  to be miniaturized to fit within a sonobouy tube, it also absorbs the forces of launch with a Charge Actuated Device (CAD) from an aircraft such as the Navy&#39;s P-3 maritime surveillance aircraft. 
     In one embodiment suggested in  FIG.  21 A , disposable launch canister  150  is provided to protect the aerodynamic surfaces of rotary wing vehicle  1  as it is launched from an aircraft traveling 150-250 knots at an altitude of 10,000 to 20,000 feet. A parachute (not shown) attached to canister  150  slows and stabilizes the descent of canister  150  which separates from rotary wing vehicle  1  at a lower altitude. Illustratively, rotary wing vehicle  1  is shown to scale and has a body length  30  of about 24 inches (51 cm), upper diameter  31  of about 2.25 inches (5.7 cm), upper rotor diameter  32  of about 28 inches (71 cm) and lower rotor diameter  33  of about 24 inches (61 cm) or less. Booster module  8  has a length  34  of about 12 inches (30 cm). First rotor  3  and second rotor  5  rotate at about 1400 RPM in hovering flight and at about or above 2000 RPM during vertical ascent and high-speed maneuvers. 
     Another embodiment contemplated by this disclosure is adapted for use with a munition for assessing target damage done by the munition. As shown in  FIG.  23   , vehicle  1  is adapted for use with the munition, illustratively shown in the drawings as a gravity-delivered bomb  160 . Bomb  160  is dropped from a launch platform such as an aircraft. In operation, gravity-delivered bomb  160  transports vehicle  1  to the vicinity of a target site whereupon vehicle  1  is released to fall away from bomb  160 , illustratively slowed by use of an auxiliary drag chute  162 , or ejected from bomb  160  by an explosive charge-actuated device, before bomb  160  reaches its target. Vehicle  1  then orbits or hovers in the target area near the impact site to observe bomb damage and transmits video and other information to a remote operator (not shown). This method of munition damage assessment is a feature of the disclosure which provides immediate battle damage assessments without requiring a launch platform to remain in the strike zone and reduces the need for subsequent strikes against the same target while minimizing risk to human crew members. 
     As shown in  FIG.  26   , motors  54 ,  61  are positioned to lie between rotor blades  20 ,  22 . Servo actuators  58 ,  59  are positioned to lie in spaced-apart relation to locate rotor blades  20 ,  22  therebetween. 
     In another illustrative embodiment motors  54 ,  61  are located below rotor blades  22  and rotating torque tube  254  runs inside non-rotating mast tube  253  for transmitting power to rotor  22  as shown, for example, in  FIGS.  28 - 31   . In another embodiment a gas engine (not shown) may be provided to generate electric power from a heavy fuel such as diesel fuel or JP8 to operate motors  54 ,  61 . In yet another embodiment, a gas engine (not shown) may be connected to torque tube  254  and rotor mount  100  through a gearbox (not shown) to drive rotors blades  20 ,  22 , also called rotors  20 ,  22 , about common axis  7 , also called rotor axis  7 . 
     Torque tube  254  may be connected directly to upper rotor hub  270  as suggested in  FIGS.  28  and  29    or to a belt or gear powered transmission and speed reduction system  271  provided at the upper end of mast tube  253  as suggested in  FIGS.  29  and  30   . Speed reduction system  271 , also called transmission system  271 , may be located at the upper end of mast tube  253  so that torque tube  254  may be configured for high-speed, low-torque operation. As a result, torque tube  254  may be of lower weight construction than a comparably sized main rotor shaft for a helicopter that must support the full flight loads of rotor hub  270  and upper rotor blades  20 . 
     Referring to  FIGS.  27 - 31   , rotary wing vehicles  250 ,  251  contemplated by this disclosure include a streamlined body  260  and other features suitable for high-speed horizontal flight. Body  260  may be adapted in some embodiments to carry one or more human pilots or one or more passenger. Rotary wing vehicles  250 ,  251  include counter-rotating rotor blades  20 ,  22  rotatable about common axis  7 , landing gear  261 , streamlined mast shroud  257 , pusher propeller  258 , and stabilizing tail fins  259 . Mast shroud  257  is generally airfoiled in cross section when viewed from above to reduce frontal drag. Mast shroud  257  is shown secured to body shell  11  and hence by screws  277  to body shell standoffs  86 ,  87 ,  88  which secure mast shroud  257  to mast tube  253  and prevent mast shroud  257  from rotating about common axis  7 . 
     As described in  FIGS.  28  and  29   , a rotor module  264  includes upper rotor blades  20 , lower rotor blades  22 , rotor control assembly  255 , rotor drive assembly  262 , and mast assembly  252 . Rotor control assembly  255  includes swashplates  56 ′,  57 ′, servos  58 ,  59 , and pitch links  125 ,  126 . Rotor drive assembly  262  includes motors  54 ,  61  with associated drive gears for driving rotors  20 ,  22  about rotor axis  7 . 
     Mast assembly  252  includes torque tube  254  running inside mast tube  253  and supported by upper mast bearing  273  and lower mast bearing  274  as shown in  FIG.  32   . Mast assembly  252  is secured to body  260  by mast brackets  266 ,  267  and mast bolts  202 . 
     Torque tube  254  is smaller in diameter than mast tube  253  leaving an annular space  275  running through the interior of mast tube  253  to act as a conduit for electrical wiring to servos  58 ,  59  and other electrical/electronic components. Wire slots  265 ,  269  are provided as entry and exits points for wiring, plumbing, and linkages (not shown). In one embodiment mast tube  253  is constructed of carbon fiber composite material and supports lateral flight loads produced by rotor blades  20 ,  22  and damps in-flight vibration of torque tube  254  especially at upper mast bearing  273 . Torque tube  254  may be constructed from carbon fiber, aluminum, or steel and may support vertical flight loads in addition to torsion. Mast bearing  273 ,  274  may be configured to support axial as well as radial loads. Because mast tube  253  is generally rigid and non-rotating, mast assembly  252  may be stronger and produce less vibration than a rotor shaft on a conventional coaxial rotor helicopter which is generally unsupported by airframe structure above the lower rotor. 
     Referring now to  FIGS.  33 - 36   , a rotor control assembly  282  in accordance with one embodiment of the current disclosure includes upper swashplate  279 , lower swashplate  280 , servo actuators  284 ,  285 ,  286 , servo ring mounts  288 ,  289  and three blade pitch Z-links  291 . While Z-link  291  may be constructed as a single piece, it is shown in the drawings as an assembly of parts consisting of a generally rigid Z-link body  292  made of glass-filled nylon and two wear-resistant universal ball links  293 ,  294  made of a softer material such as unfilled nylon. Universal ball links  293 ,  294  fit into link recesses  299 ,  300  in Z-link body  292  and are attached by screws  295 . 
     Simultaneous, uniform, axial displacement of all three Z-links  291  in rotor control assembly  282 , also called swashplate control assembly  282 , parallel to common axis  7  causes swashplate  279  and swashplate  280  to move axially along common axis  7  which displaces pitch links  119  thereby changing the collective pitch of rotor blades  20 ,  22  simultaneously. Non-uniform and independent axial displacement of Z-links  291  causes swashplates  279 ,  280  to tilt simultaneously inducing a cyclic pitch control in rotor blades  20 ,  22 . Z-links  291  are also constrained to move parallel to common axis  7  by anti-rotation tabs  287  appended to ring mounts  288 ,  298  and act as swashplate anti-rotation links. 
     Z-link body  292  is configured to hold universal ball links  293 ,  294  at a fixed differential phase angle  290  so that non-uniform axial displacement of Z-links  291  parallel to common axis  7  in direction  298  causes swashplate  279  and swashplate  280  to tilt in different directions which affects the relative cyclic phase angle of rotor blades  20  and  22 . Differential phase angle  290  is shown as 90 degrees but may lie between about 60 to about 120 degrees depending on the characteristics of rotor blades  20 ,  22  and their speed of rotation. Differential phase angle  290  may be changed by varying the length of universal ball links  293 ,  294 . 
     Z-link  291  aligns the cyclic phase angles of upper rotor blades  20  and lower rotor blades  22 . Rotor phase angle can be described as the angle measured between the cyclic pitch control input of a swashplate to a rotor system of rotating rotor blades and the resulting flapping motion of the rotor blades and apparent tilt of the rotor disk. Normally the phase angle of a single rotor helicopter is close to 90 degrees. 
     Because of the aerodynamic interaction of the upper and lower blades on a coaxial rotor helicopter, however, the rotor phase response of each rotor on a coaxial rotor helicopter is much different than 90 degrees. For instance as illustrated in  FIG.  37   , if upper swashplate  279  and lower swashplate  280  are tilted forward in direction  297 , upper rotor blades  20  will appear to tilt in upper rotor phase direction  302  and lower rotor blades  22  will appear to tilt in lower rotor phase direction  303  which means that the absolute upper and lower rotor phase angles are each about 45 degrees. The phase angle difference  304  therefore is about 90 degrees. When upper swashplate  279  and lower swashplate  280  are each rotated 45 degrees about common axis  7  by the fixed differential phase angle  290  of Z-links  291  before being tilted then upper rotor blades  20  and lower rotor blades  22  will both appear to tilt in direction  297 . At this point upper rotor blades  20  and lower rotor blades  22  are said to be in phase with each other. Rotors that react in phase with each other produce powerful control forces. 
     As illustrated in  FIGS.  38  and  39   , a rotary wing vehicle according to the current disclosure includes a streamlined fuselage or body  260 , a rotorcraft power and control system  306 , a co-axial, counter-rotating rotor system  307  capable of producing vertical lift and a rearward facing propeller  258  capable of producing horizontal thrust. 
     In operation, power from a motor or engine  309  turns first stage pinion gear  311  which turns crown gear  312 , 313  in opposite directions as described in  FIGS.  38  and  39   . Crown gear  312  is connected by a transfer shaft to second stage pinion  314  which drives lower rotor main gear  316  and lower rotors  22 . Crown gear  313  is connected by a transfer shaft to second stage pinion  315  which drives upper rotor main gear  317 , torque tube  254  inside mast  319  and upper rotors  20 . A belt drive system consisting of pulleys  321 , 322  and V-belt  323  drive propeller shaft  324  from the aft end of motor  309 . 
     As illustrated in  FIG.  40   , a non-rotating structural mast  319  according to the current disclosure is configured with interior passageways or conduits  325  to accommodate both mechanical and electrical power and signal transmission components. Mast  319  may include center column  326  and outer sheath  327  which are generally circular in cross section and connected by radially extending ribs  328  which function to both separate and stiffen center column  326  and outer sheath  327 . In operation torque tube  254  runs between bearings  273 ,  274  (see  FIG.  32   ) inside center column  326  to transmit rotary motion from a power source located below mast  319  to rotor blades  20  located near the upper end  318  of mast  319 . Bearings  273 ,  274  act to align mast inside of center column  326  and prevent torque tube  254  from bending or touching the interior surface of center column  326 . Torque tube  254  is mechanically separated from wiring, plumbing, hoses and linkages (not shown) which are located between center column  326  and outer sheath  327  in interior conduits  325 . In essence, center column  326 , outer sheath  327  and ribs  328  form a plurality of signal and power conduits which effectively separate mechanical, electrical and fluidic power and signal lines running inside mast  319 . 
     Referring now to  FIG.  41 - 43   , a non-rotating structural mast  330  according to the current disclosure is configured with six interior passageways  331  to accommodate swashplate linkages  332  that transfer mechanical control signals from servos actuators (not shown) located below lower rotors  22  to swashplates  279 ,  280 . Mast  330  may include center column  333  and outer sheath  334  which may be generally circular in cross section and connected by radially extending ribs  335  which function to both separate and stiffen center column  333  and outer sheath  334 . In operation torque tube  254  runs inside center column  326  to transmit rotary motion from a power source located below rotor blades  22  to rotor blades  20  located near the upper end  336  of mast  330 . 
     Apertures or slots  342  may be provided in outer sheath  334  to accommodate entry and exit of wiring, plumbing, hoses (not shown) and swashplate linkages  332 . A feature of the current disclosure is that ribs  335  and center column  33  act to transmit structural loads around apertures  342  thereby improving the structural integrity of mast  330  especially when many power and signal lines are routed through mast  330  and much of outer sheath  334  is perforated by slots or holes. Another feature is that apertures  342  may extend completely to an end  337  of mast  330  to allow removal of mast  330  from an aircraft during maintenance operations. In one embodiment, power and signal lines running inside mast  330  may be removed and reinstalled without first removing plugs and connectors that may not easily fit through interior passageways  331  thereby reducing maintenance costs. Yet another feature of the current disclosure is that mast  330  may be economically manufactured, for instance, in an extrusion process from aluminum alloy 7075 or in a pulltrusion process from epoxy impregnated carbon fibers for low weight and high strength. 
     As shown in  FIGS.  44 A and  44 B , each swashplate linkage  332  may be assembled from lower slider  338 , upper slider  339 , slider pushrod  340  and pitch control link  341 . Lower sliders  338  may be connected to a servo actuator (not shown) to move swashplate linkages  332  axially inside interior passageway  331  of mast  330 . Upper sliders  339  are pivotably connected to pitch control links  341  which transmits axial motion of swashplate linkages  332 , also called swashplate sliders  332 , to swasplates  279 , 280 . Slider pushrod  340  is shown with threaded ends and rigidly connects upper slider  339  and lower slider  338  to move as a unit. 
     Three servo actuators (not shown) connected to lower sliders  338  may cooperate to move three swashplate linkages  332  to control upper swashplate  279  and the cyclic and collective pitch of rotor blades  20 . Three additional servo actuators (not shown) connected to lower sliders  338  may cooperate to move three swashplate linkages  332  to control lower swashplate  280  and the cyclic and collective pitch of rotor blades  22 . While shown in the drawings with pitch control link  341 , swashplate linkages  332  may also incorporate Z-link  291  in place of pitch control link  341  in which case only three servos would be needed to control the cyclic and collective pitch of both rotor blades  20 ,  22 . 
     As illustrated in  FIGS.  45  and  46   , a rotary wing vehicle  350  in accordance with the present disclosure includes a streamlined fuselage or body  351 , a co-axial, counter-rotating rotor system with counter-rotating rotor blades (not shown) capable of producing vertical lift and a rearward facing propeller  353  capable of producing horizontal thrust. A non-rotating backbone or mast  330  supports a plurality of rotary output servo actuators  354  located behind mast  330  and a plurality of rotary output servo actuators  355  located in front of mast  330 . Servo actuators  354 ,  355  are configured to lie in close proximity to a longitudinally extending plane defined by common axis  7  and longitudinal axis  356  to reduce the forward-facing surface area of the servo actuators  354 ,  355  in high-speed forward flight. This reduces the width of a shroud (not shown but similar to shroud  257  in  FIG.  27    and shroud  368  shown in  FIG.  48   ) needed to cover servo actuators  354 ,  355  and minimize aerodynamic drag in high speed forward flight. Bolt holes  357 , as shown in  FIG.  46   , are provided to mount a streamlined mast shroud such as shroud  257 . One feature of the current disclosure is that control system components such as servo actuators  354 ,  355  are located in front of and behind mast  330  to minimize the width of the mast assembly to reduce drag in forward flight. 
     Another embodiment of a rotary wing vehicle  360  is shown, for example in  FIGS.  47 - 57   . Rotary wing vehicle  360  includes a streamlined fuselage or body  361 , a co-axial, counter-rotating rotor system with counter-rotating rotor blades  362 ,  375  capable of producing vertical lift and a rearward facing propeller  353  capable of producing horizontal thrust. A non-rotating mast  364  supports mast sleeve  366  and a plurality of linear (screwtype) servo actuators  365 . In one example, the linear (screwtype) servo actuators  365  may be Moog model 880 Electric Linear Servo Actuators that are mounted thereto by brackets or arms protruding therefrom. Servo actuators  365  are configured to lie in close proximity to a longitudinally extending plane defined by common axis  7  and longitudinal axis  367  to reduce the width and aerodynamic drag of mast shroud  368  in high-speed forward flight. Engine  363 , which may be a GE T700 turboshaft engine for example, is provided to turn upper rotor  362  about common axis  7  through gearbox  369 , upper rotor drive gear  370  and upper rotor torque tube  379 , and to turn lower rotor  375  through gearbox  369  and lower rotor drive gear  371  attached to lower rotor shaft  380 . 
     A feature of the current disclosure is that non-rotating mast  364  may support aircraft components inside of mast shroud  368  to take advantage of the air wake produced by mast shroud  368  in high-speed forward flight. Electronic or hydraulic components  372 , including, for example, hydraulic motors and hydraulic valves, and antennae  373  may be supported by non-rotating bracket  374  in some embodiments. This reduces the need for space inside the body  361 , also called fuselage  361 , of rotary wing vehicle  360  and places electronic or hydraulic components closer to servo actuators  365 . 
     Non-rotating mast  364  may be fabricated from a metal or carbon fiber composite material and include channels  376  extending axially along an exterior surface of mast  364  to accommodate electrical bus inlays  378  as suggested in  FIGS.  50 - 52   . Electrical bus inlays  378  extends from a point  390  between upper and lower rotors  362 ,  375  to a point  391  below the lower rotor  375  and between upper rotor drive gear  370  and lower rotor drive gear  371  to facilitate transmission of electrical and/or hydraulic power and signals from components located in fuselage  361  of rotary wing vehicle  360  to other components located between upper rotor  362  and lower rotor  375  or above the upper rotor  362 . Electrical bus inlays  378  may include a protective sheath made of a non-conducting material such as silicone and contain a plurality of copper conductors or hoses  382 . In one embodiment mast sleeve  366  slides over mast  364  to provide a mounting structure for servo actuators  365  and bracket  374  and a smooth exterior running surface for swashplates  384 , 385 . Apertures  387  may be provided in mast sleeve  366  to provide access to copper conductors or hoses  382  for electrical or hydraulic connections (not shown) to other components such as servo actuators  365  and flight control system electronics (not shown). In operation a plurality of electrical wires and/or hydraulic hoses (not show for clarity) may connect to bus inlays  378  at copper conductors or hoses  382  to transmit electrical or hydraulic power and signals to and from other control system components such as a flight management system computer (not shown), servo drivers (not shown), hydraulic motor  372 , hydraulic values (not shown), and generators (not shown). A sturdy truss structure  388  may be provided to connect mast  364  to fuselage  361  of rotary wing vehicle  360 . 
     An important feature of the current disclosure is the reduction of aerodynamic drag in high-speed flight. To reduce the width and associated drag of mast shroud  368 , swashplates  384  and  385  are configured to locate all six servo actuators  365  in close proximity to a longitudinally extending plane defined by common axis  7  and longitudinal axis  367  as illustrated in  FIG.  53   . Swashplate arms  392  and  393  are closer to each other than arms  393  and  394 . As shown in  FIG.  54   , angle  395  is about 90 degrees or less. Swashplates  384  and  385  are also rotated 180 degrees relative to each other about common axis  7  so that servo actuators  365  may be interleaved around the circumference of mast sleeve  366  for a very compact installation. 
     One feature of the disclosure is the non-rotating hollow core tube  40 , mast  330 ,  364  or cruciform beam structural backbone that can, in some embodiments, double as a conduit for wiring and plumbing. A method or system of assembling mechanical and electrical components to the core or backbone is described to promote ease of assembly of a variety of aircraft from a kit of basic modules. 
     Another feature is that each of the rotors  20 ,  22  of the coaxial system of the current disclosure are driven by one or more separate electric motors, and the motors are positioned to lie on opposites sides of the rotors, with power transmission to and between the motors accomplished through electrical wiring (passing through the hollow core) instead of mechanical shafting, clutches, and gears. Compact rotor assemblies support the rotors for rotation without the need for traditional rotating coaxial shafting. 
     Still another feature is that a swashplate control system and one or more electric motors may be provided for each rotor and may be positioned to lie on opposite sides of each rotor thereby simplifying the mechanical and electrical connections needed to drive and control the rotors. Rotor modules are provided to quickly and easily assemble systems of rotors to the hollow core. Multiple rotor modules and swashplates are controlled by a single group of servos housed in a module. 
     Another feature of the disclosure is the provision of phase links to produce differential phase control of the upper and lower rotors simultaneously. In some embodiments, fixed-phase links can provide collective and cyclic control of both rotors with only three rotor control servos instead of the four to six servos generally required for coaxial rotor control. 
     Another feature is that full collective and cyclic control of the upper and lower rotor blades of a coaxial helicopter can be accomplished with servo actuators located below the lower rotor so that the axial distance between the upper and lower blades can be minimized. 
     Another feature is that a streamlined, non-rotating body shell may be mounted between the upper and lower rotor blades of a coaxial helicopter to reduce drag in high-speed forward flight. 
     Yet another feature of one embodiment is that power and control signals may be passed from a point located below the lower rotor blades to a point located between the rotor blades to facilitate locating the rotor control system, radio electronics, antennae, and other electrical and control system components between the rotor blades to make productive use of the space between and the blades in high speed forward flight. 
     Yet another feature of one embodiment is that upper rotor blades  20  may be driven by a torque tube  254  running inside the mast tube  253  and connected to a motor  54  or engine located below rotor blades  22 . Both upper and lower rotors may be driven by a single gas-powered engine located below the rotors if desired. 
     An additional feature is that folding rotor blades  148 ,  149  are of unequal length. On the current disclosure with counter-rotating rotors  3 ,  5 , folding blades  148 ,  149  of unequal length reduce the chance that the blades will contact one another as they fold at high speed during a crash-landing. 
     Another feature is that a mounting structure is provided between counter-rotating rotors  20 ,  22  to support a body shell  11  or other type of aerodynamic fairing between rotor blades  20 ,  22 . Body shell  11  protects the control assembly  255  from weather and reduces the air resistance of exposed servos  58 ,  59 , swashplates  56 ′,  57 ′, and pitch links  125 ,  126 , also called pushrods  125 ,  126 . 
     Another feature of the disclosure is a method of improving energy and power density on UAV&#39;s which can include a booster module  8  which is separable from the main vehicle in flight. A booster module  8  is provided to operate the UAV during a first flight phase. At the end of the first flight phase, the booster module falls away thereby reducing the weight of the UAV for continued operation in a second flight phase. On electric powered UAV&#39;s, the power module may comprise a pack of batteries with or without an auxiliary lifting surface which is jettisoned in flight after the battery power is depleted, or payloads specific to a particular mission.