Abstract:
An air vehicle  1  has a port and starboard wing  2  capable of transitioning in mid-flight from fixed to counter-rotation. This generates lift at low airspeeds and enables a short landing to be performed without the need for a runway. The counter rotating wings are intermeshed at 90 degrees and angled slightly to avoid collision. Deployable counterweights  5  balance the rotating wings.

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
       [0001]    This application claims the right of priority of the UK application no. GB1103439.4 filed on 1 Mar. 2011 by the inventor. 
       FEDERALLY SPONSORED RESEARCH 
       [0002]    Not Applicable 
       SEQUENCE LISTING OR PROGRAM 
       [0003]    Not Applicable 
       BACKGROUND OF INVENTION 
       [0004]    Under certain circumstances the desired destination for an air vehicle lacks a suitable runway. In these situations an air vehicle capable of vertical or very short landing would be an advantage. 
         [0005]    Helicopters are capable of vertical take-off and landing. Helicopters use rotary wings powered by an engine through a gearbox. This results in a complex system limiting useful load, and range and increasing costs when compared to a fixed wing airplane. 
         [0006]    An auto-gyro also uses rotary wings but unlike a helicopter the rotors are powered by external upward airflow. The upward airflow is generated by either a slow descent of the auto-gyro or by forward flight with the rotors tilted backwards which in this case requires a form of the propulsion. An auto-gyro can land gently and almost vertically but with a small amount of forward speed still present. 
         [0007]    The helicopter and auto-gyro have a limited maximum forward speed due to the asymmetric lift between a retreating blade and an advancing blade. A rotor also generates high aerodynamic drag as compared to an equivalent fixed wing aircraft. 
         [0008]    An air vehicle that combines the advantages of a fixed wing for cruise with a rotary wing for landing provides an efficient means of air transport for the following (but not limited to) types of aircraft:
   I. A military cargo glider where after landing it is discarded.   II. A Navy carrier aircraft where (especially when unmanned) a gentle landing is required but could be reconfigured to take-off using the catapult system.   
 
         [0011]    The Herrick convertible aircraft comprised an upper wing on top of a mast that was locked into position during cruise and when combined with the lower wing operated as a biplane. When the wing was unlocked the top wing was free to rotate as an auto-gyro, reference U.S. Pat. No. 2,699,299. 
         [0012]    The rotor was initially powered by rubber bungee cords contained inside each upper wing half, running through an aluminum tube. Prior to take-off, two people rotated the wing twice in the opposite direction to the auto-rotation. The upper wing was then locked in the biplane cantilever wing position. When the pilot released lock during the flight, the bungee cords rotated the wing spun twice at 60 revolutions per minute. The now-spinning wing rotated freely and increased its speed to 220 revolutions per minute in auto-rotation. 
         [0013]    This type of approach requires the airflow over one half of the wing being reversed during rotation. To accommodate this bi-directional airflow, the wing section shape is bi-convex or somewhat lenticular. This shape is widely used for supersonic aircraft and missiles but grossly inefficient when used at subsonic speeds. 
         [0014]    The following invention allows fixed wing to convert to auto gyrating wings but without the disadvantages of using bi-convex wing shapes. 
       SUMMARY 
       [0015]    It is proposed that an air vehicle comprising a port and starboard wing capable of transitioning in mid-flight from fixed to rotation in opposite directions to generate lift at low airspeed thus enabling a short landing without the need for a runway. 
         [0016]    For each rotating wing, a counter weight ensures that the rotation is balanced. The counterweights are stored within the wing during normal flight and deployed in the plane of the rotor during transition. The act of deploying the counter weight spins the rotor from its static position to a rotor speed where the airflow begins to turn the rotor and provide lift. 
         [0017]    Both rotors turn in opposite directions. This requires the rotors to be inter-meshed at 90 degrees and set at a slight angle to avoid collision. 
         [0018]    The embodiment chosen is based on a military cargo glider simply because the glider is simpler allowing the drawings to show the invention in better clarity. 
         [0019]    Many changes, modifications and substitutions may be made to the following embodiment without departing from the spirit and scope of the invention. The invention is not limited by the following embodiments and their descriptions. 
       Military Glider 
       [0020]    Military operations may require the rapid deployment of a bulky cargo such as armored fighting vehicles (hereafter abbreviated to AFV) over great distances to landing areas that are remote from the sea. In such circumstances deployment by air would be required. 
         [0021]    Ideally, military cargo planes would land and then disembark the cargo. However, with no available runway or unfavorable terrain, landing may be impossible. 
         [0022]    Cargo pallets and small vehicles can be parachuted from the back of a military cargo plane in flight. Modern large ‘wing-type’ parachutes accommodate significant loads and are steerable but are still limited to the smallest of AFV. 
         [0023]    Most military cargo aircraft can only transport a single AFV in a single flight. Therefore a substantial number of flights would be required in a short period to transport a significant number of AFVs. This requires substantial number of military cargo aircraft which would be costly especially when such deployments are rarely needed. 
         [0024]    Although military gliders have given way to modern military transport planes and helicopters, a glider still provides unique advantages. 
         [0025]    A military glider can be to be manufactured for the fraction of the cost of a military transport aircraft especially when designed for a single operation and to be flown unmanned using modern avionics. A military glider also lacks engines and fuel tanks which further simplifies design and lowers costs. 
         [0026]    A military glider can be towed by numerous types of aircraft such as commercial airliners. The tow aircraft could possibly be lightly modified at a typical maintenance site. 
         [0027]    However, a military glider still requires a suitable area to make a standard long landing. Alternatively, retro rockets, drogue parachutes, and other devices provide the means of shortening the landing (arrestor cable for carriers). The abrupt landing results in significant forces that put a large strain on the airframe and on the AFV and may result in equipment damage. 
         [0028]    A military glider that transitions to an auto-gyro by converting the fixed wings into counter-rotary wings would be an advantage with regards to soft landings. This embodiment makes reference to such a glider using the following drawings. 
     
    
     
       DRAWINGS 
         [0029]      FIG. 1  is an isometric view of the glider. 
           [0030]      FIG. 2  is a side view of the glider at take off. 
           [0031]      FIG. 3  is a side view of the glider after landing and debarkation. 
           [0032]      FIG. 4  is a schematic view of the shaft layout for both the port and starboard wing and their inter-connection. 
           [0033]      FIG. 5  is a schematic view of the counter weight deployment system and boundary layer lift system. 
           [0034]      FIG. 6   a  is a physical view of the counter weight deployment system and boundary layer control lift augmentation system 
           [0035]      FIG. 6   b  is a section drawing of the rotor mounting hub and counter weight deployment gears. 
           [0036]      FIG. 6   c  is a section drawing of the piston and pressure regulator used to deploy the counterweight. 
           [0037]      FIG. 6   d  is a section drawing of the pressure tube used to deploy the counter weight at the end of piston travel. 
           [0038]      FIG. 6   e  is a cross section drawing of the counter weight shaft, 
           [0039]      FIG. 7   a  is a section drawing of the synchronizing gearbox. 
           [0040]      FIG. 7   b  is a side view of the logarithmic gears at the start (a) and at the end (b) of the transition period 
           [0041]      FIG. 7   c  is a side view of the logarithmic gears at the start (a) and at the end (b) of the transition period 
       
    
    
     DETAILED DESCRIPTION 
       [0042]      FIG. 1  shows the glider  1  with its two wings  2  in the high position above the fuselage  3 . The wings have a long aspect ratio and a thin chord with a standard airfoil shape for efficient use in both fixed and rotary modes. The wings are shown in the fixed mode with a secondary dashed outline of the starboard wing shown 90 degrees to the port wing as would be the case when rotating. 
         [0043]    Two struts  4  extend upwards from the fuselage  3  and are firmly attached to each wing  2 . At the wing roots, two counter weights  5  are shown in their non-deployed position (but shown fully deployed for the dashed starboard wing). The wing incorporates slot  6  that runs along on the top of the wing and is part of the boundary layer control system. 
         [0044]    The tail includes two fully actuated fins  8  in a ‘inverted V’ and a pitch rocket  9 . Also a turbine  7  powers the on-board systems. 
         [0045]    The front of the fuselage incorporates the nose cone  10  with a tow line  11  that extends to the tow plane. 
         [0046]    The long aspect ratio wing reduces drag. The thin airfoil section allows for high speed flight near the transonic region without the need for sweep back. Also the thin airfoil section promotes low-drag laminar flow. 
         [0047]    The struts mechanically support the long thin wings whilst minimizing structural weight. The struts also lock the wings into the fixed position during the flight. The transition process from fixed to rotary begins when the strut is commanded to detach from the wing. 
         [0048]    The struts act as temporary wings during the transition period where the rotating wings have not gained sufficient speed to provide lift. Although not sufficient to provide level flight, the struts provide stability to the glider through this phase. 
         [0049]    The rear fins provide yaw and roll authority to the glider. The fins also provide roll through asymmetric lift caused by the yaw rotation. Furthermore with the fins being below the gliders center of gravity a roll moment is also generated. 
         [0050]    The tail is long to allow the fins and the pitch rocket to have a robust momentum arm. 
         [0051]    The large wingspan enables the wings to provide sufficient lift at low rotational speeds. An estimated rotation speed of less than 80 rpm has been calculated which is significantly less than traditional auto-gyros and helicopters rotor speed of over 300 rpm. The low rotational speed keeps the centrifugal loads within tolerable limits for the proposed wing structure but sufficient to provide the required stiffness for the wing to behave as an efficient rotor disc. 
         [0052]    The auto gyrating wing experiences three separate flow regimes along its length. This is due to the radial change in velocity when combined with the airflow velocity passing through the rotor. At the root of the blade, the combined air velocity results in an angle of attack too high resulting in blade stall. A little further out from the hub the angle of attack results in a resultant force which not only provides lift but also a forward acceleration to the rotor. At the tip, the angle of attack results in a force which again provides lift but this time a braking force to the rotor. In steady flight, the accelerating and breaking forces cancel leaving the wing to spin at a steady speed. 
         [0053]    A thin wing section exhibits a maximum lift coefficient less than a standard thicker wing section. As such the stall region during rotation, near the wing root, is more pronounced than would otherwise be. This small loss of lift can easily be re-claimed using boundary layer control and when applied separately to each wing provides roll control during the rotor wing phase. 
         [0054]    The slot  6  incorporated in the wing sucks low energy air from the stall region. Energetic air from the main airflow fills the void and allows the flow to ‘un-stall’ and re-attach to the wing increasing the overall lift of the wing. 
         [0055]    The glider proposed here does not comprise pitch or cyclic control but only rigid wings simplifying the design. 
         [0056]    With a twin rotor system that has a side by side configuration, the roll moment caused by the asymmetric lift resulting from forward flight is canceled out. High forward speeds during the rotary phase are not expected because transition takes place near the landing site. 
         [0057]    As the transition occurs at low speed near the landing site, the pitch control is not warranted. Should an inadvertent pitch down moment occur and the glider picks up speed, the fins  8 , now loaded by the airflow, would pitch the glider in the traverse axis to reduce the speed. 
         [0058]    However a pitch up provides a significant increase in lift during final landing phase. Therefore a small rocket motor  9  positioned in the tail section ignites just before landing. This pitches the glider upwards producing additional lift by using the kinetic energy in the spinning rotors and in the counter weights. 
         [0059]    The boundary layer system  6  when used on either the port or starboard wing provides a roll moment by asymmetric lift between the two wings. A yaw moment in the direction of roll moment is also generated as the two wings are mechanically linked. The boundary layer control when activated results in an increase in rotor speed by the mechanism stated earlier. A breaking torque for one wing and a driving torque for the other wing are set up. The resulting torque reaction occurs in the same direction as the wings are counter-rotating causing the glider to yaw. In a similar way the glider in auto gyration mode would have the tendency to ‘weathercock’ even without the influence of the fins  8 . 
         [0060]    The counter rotation of the two wings generates a slight rocking motion diagonal to the body axis but this would be short lived and non-troublesome for unmanned use. 
         [0061]      FIG. 2  shows the glider at the point of take off. A jettisonable undercarriage  12  detaches from the glider at take off. This removes the need for a retractable undercarriage. 
         [0062]      FIG. 3  shows the glider after landing. Embedded detonating chords shatter the nose cone to allow the AFV  13  to disembark from the glider. This removes the need for an actuation system and locking mechanism allowing standard bolts to attach the nose cone to the fuselage after the AFV has been loaded prior to take-off. 
         [0063]    Without the undercarriage the underside of the fuselage crumples on impact protecting the load during a hard landing. 
         [0064]      FIG. 4  shows the mechanical layout of the port and starboard wing.  FIG. 4  also shows the wing dihedral in relationship to the fuselage. In this particular design, the dihedral is 5 degrees which allows the rotating wing  2  to miss the hub  14  of the adjacent wing. The wings shown are at 90 degrees to each other. A geared hub  14  attaches the wing to the glider. The counter weights  5  are shown in their extended position. A drive shaft protrudes from each hub to bevel gears  15  that turn the drive perpendicular towards the synchronizing gearbox  16 . This arrangement ensures the wings inter-mesh and avoid hitting each other. 
         [0065]    Counter weights  5  opposite each wing cancel the centripetal loads that would cause premature failure to the hub  14 . Owing to the size of the wings, the counter weight needs to be at the end of long extension arm  17 . To avoid excessive drag during the flight, the extension arm is stored inside the wing and deployed when the glider transitions to rotary wing. 
         [0066]      FIG. 5  shows a schematic layout of the counterweight deployment system  30 . The pressure cylinder  37  sits inside the wing root and is structurally part of the wing spar. Housed within the cylinder  37  is the counter weight extension arm  17 . The counterweight  5  also acts as a pressure vessel storing compressed air that extends down the length of the extension arm  17  to a high pressure regulator  31 . A firing spring  32  sits between the extension arm piston and the cylinder end face and keeps the pressure regulator closed during the flight. The extension arm comprises a rack tooth profile that engages the gears of the rotor hub  14 . 
         [0067]    When the wing detaches from the strut, the firing spring pushes on the extension arm  17  allowing the regulator  31  to pressurize the cylinder  37 . The pressure forced the extension arm outwards deploying the counter weight. As the counter weight is deployed the extension arm also drives the gear hub  14  which spins the wing as the annulus of the gear hub  14  is fixed to the fuselage. The wing spins to a suitable speed where aerodynamic auto-rotation begins. 
         [0068]    The centripetal force that begins to act on the counter weight and hence further the spin acceleration through the hub  14  is canceled by the opposite force generated by the Coriolis Effect. 
         [0069]      FIG. 5  also shows a schematic layout of the boundary layer control system  29  with a low pressure regulator  33 , solenoid valve  34  and venturi tube assembly  35  attached to the wing slots  6 . A rotary transformer  36  connects the solenoid valve to the cabling inside the fuselage. 
         [0070]      FIG. 6   a  shows the physical layout of both the counterweight deployment system and the boundary layer control system. These two sub-systems are combined into one physical layout. 
         [0071]    The pressure cylinder  37  fits inside the wing and is part of the spar. The counter weight  5  incorporates two solid masses  38  made from depleted uranium or tungsten which provide the majority of the counter weight mass to the composite built wing. Sandwiched between the two masses is a high pressure reservoir  39  that provides compressed gas for the counter weight deployment and the boundary layer control system. 
         [0072]    The hub  14  is based on an epicyclic gearbox, but here the extension arm replaces the sun gear. The extension arm drives the planetary gears inside the annulus. The annulus remains fixed to the fuselage with the planetary gears rotating the wing. 
         [0073]    This arrangement allows the pressure cylinder  37  and hence the wing spar to be in line with the rotation axis of the hub and maximizes the pressure cylinder diameter inside the thin airfoil section. Also the extension arm can better match the wing rotation in displacement and how pressure relates to acceleration. 
         [0074]      FIG. 6   b  shows a cross section of the hub  14 . The top housing plate  40  is attached to the pressure cylinder  37  through a bolted flange  41 . The top housing is bolted to the bottom plate  42  through three spindles  43 . Two of the spindles mount the planetary gears  44  through tapered bearings  45 . The third spindle, not shown in this cross section, mounts a guide wheel (ref  FIG. 6   a ) that ensures correct alignment and gear meshing of the extension arm  17  through the hub  14 . 
         [0075]    A long pin  46  attaches the bottom plate  42  to the output shaft  47 . A bevel gear  48  at the end of the output shaft meshes with a perpendicular shaft  28  that allows the transmission straight to the synchronize gearbox. The bottom plate  42  sits inside the fixed annulus ring  49  through two sets of bearings. The larger thrust bearing  50  supports entire load of the glider and is therefore substantial. The smaller tapered bearing  51  is a taper bearing provides lateral support any unbalance loads and negative g loads although these are expected to be small. The extension arm  17  passes through the hub to the counterweight not shown. 
         [0076]    The bottom plate also includes the rotary transformer  36  which allows the electrical power from the avionic module to be transmitted across to the rotating wing rotor. The housing is not shown just the inner and outer coils for the purpose of illustration. 
         [0077]      FIG. 6   c  shows the boundary layer control system  29 . The boundary control system feeds from the pressure cylinder  37  with the remaining compressed air that has been used to deploy the counter weights. 
         [0078]    The boundary control system comprises a low pressure regulator  33 , a solenoid valve  34 , and a venturi tube assembly  35 . The low pressure regulator  33  and solenoid valve  34  are fixed to the wing spar  2  which in turn is bolted to the pressure cylinder  37 . 
         [0079]    Low pressure air switched on and off by the solenoid valve  34 , passes through a connecting pipe  55  and discharges at high velocity from the pressure nozzle  52  into a mixing tube  53  and out through the diffuser  54 . This generates a vacuum in slot  6  by means of the venturi effect and thus sucking stagnant air allowing the re-energizing air from the air stream to form an attached boundary layer. 
         [0080]    An AC coil when energized operates the solenoid valve  34 . A spring returns the valve to the off position when the electrical energy is removed. A rotary transformer  36  (reference  FIG. 6 ) allows the AC electrical power to pass from the fuselage to the rotating wing. 
         [0081]    The counter weight pressure cylinder  5  (reference  FIG. 6 ) replenishes the cylinder  37  with compressed air when air is discharged for the boundary layer control system. 
         [0082]      FIG. 6   d  details the high pressure regulator  31  plus the final cross section of the pressure cylinder  37  with its end cap  55 . The pressure regulator comprises a piston  56 , inner pressure tube  57 , and a pressure spring  58  with a washer  59 . The inner pressure tube screws directly into the extension arm  17  with the pressure spring between it and the piston. The pressure spring supported by the flange end of the extension arm pushes against the piston  56  and valve seating  60  allowing high pressure air to flow into the air cylinder. When the desired pressure is obtained in the pressure cylinder  37  the spring force is matched and shuts the valve. This regulates the air pressure. 
         [0083]    A collar  61  prevents excessive travel of the piston  56  that could block the high pressure holes  62  and eventually stops the piston from dropping of the extension arm. 
         [0084]    The outer diameter of the piston  56  and inner diameter of the pressure cylinder  37  provide a clearance fit that allows a trade off between ease of manufacture for the pressure cylinder  37  and the dynamic seals  63  to function correctly. Also the clearance fit keeps the friction between the piston  56  and the cylinder  37  to a minimum during the counter weight deployment. 
         [0085]    The clearance fit may result in a small leakage of the compressed air. Although a negligible amount of air would be lost during the short deployment period, the leakage over time would result in a complete discharge of the system. Therefore a separate firing spring  32  sits compressed between the end cap  55  and the piston  56  to ensure that the valve seat  60  pushes firmly against the valve stem  64  thereby sealing the compressed air. 
         [0086]    When the wing  2  detaches from the strut  3  (reference  FIG. 1 ), the firing spring  32  pushes the piston  56  and extension arm  17  outwards. This allows the pressure regulator  31  to begin operating and pressurize the cylinder. The compressed air metered into the cylinder at constant pressure through the regulator deploys the extension arm and counterweight. 
         [0087]    The pressure regulator  31  resides at the end of the extension and not at the exit point of the counterweight  5 . This allows the compressed gas to flow along the narrow passage way inside the extension arm at a higher pressure but with a lower flow velocity minimizing its flow restrictions. 
         [0088]      FIG. 6   e  shows a detailed section of the hub end of the pressure cylinder  37 . Here, the pressure cylinder comprises an inner sleeve  65  made from a soft material such as brass. An interference fit that exists between the inner sleeve  65  and the piston.