Abstract:
A supplemental thrust system for an airplane comprising a submerged inlet duct, a diffuser section in communication with the inlet duct, and a duct outlet in communication with the diffuser section, an outlet nozzle in communication with the duct outlet, a turbocharger unit comprising one or more turbocharges in communication with each other to provide compressed air to the engine and the cabin, a heat exchanger unit comprising a plurality of heat exchangers in the diffuser section to provide cooling liquid to cool pressurized air in the cabin, engine intake air and the engine jacket, where each turbocharger is coupled to the intake manifold of an engine to increase the power of the engine by introducing compressed air into the manifold.

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
       [0001]    This application claims priority to U.S. Provisional Patent Application Ser. No. 61/753215, filed Jan. 16, 2013, which is incorporated by reference herein in its entirety. 
     
    
     BACKGROUND OF THE INVENTION 
       [0002]    The present invention was borne out of frustration with the cost and inefficiency of the airlines&#39; hub-and-spoke transportation model. This model was conceived by the airline industry, initially in an attempt to restrain passengers from using interline transfers to arrive at their destinations. It requires dense concentrations of passengers both at the relatively few hub facilities and in ever larger aircraft flying to fewer and fewer destinations. The inefficiencies for the traveler arise out of the time wasted traveling long distances from their true origin to the large hub or major airport, enduring the lengthy lines at check-in and security check points, and the ever-longer boarding process on the ever larger aircraft. In addition, the traveler must often fly to cities that are well out of the way to his final destination, and transfer with additional wasted connection times. The result is that for short trips (approximately 500 miles) average speeds reduce to the vicinity of 100 mph, and many longer trips that involve just one connection drop to 200 to 300 mph average. This inefficiency raises costs for the consumer, especially where the inefficiencies require overnight stays in order to catch connecting flights. There is an additional factor which is a disadvantage of the current hub and spoke system. The current system creates large concentrations of people, both at terminals and in ever larger aircraft, that create prime targets for terrorist activity. Larger numbers of much smaller aircraft operating in a widely distributed transportation system would present a more difficult target for any significant military or terrorist activity. 
         [0003]    Clearly, there are compelling reasons for wanting an air transportation system that is economically superior to our current air transportation system in acquisition, operation and maintenance costs. To be a viable competitor, the system should have true origin to true destination speeds that significantly exceed current system speeds. It should require no additional infrastructure, and it should package passengers in small enough units that both the passenger load and the aircraft are militarily insignificant targets. To be truly competitive, it should provide non-stop transcontinental and intercontinental travel from any local airport to any other local airport. And ticket prices should be highly competitive with current average ticket prices. 
         [0004]    Such a transportation system requires a unique aircraft. It must be capable of operation from any current airfield. Preferably, it would have operating costs well below current costs and competitive with commercial airliners, cruise at higher system speed than current commercial aircraft, have a longer range with full passenger and luggage load than most current business aircraft, provide passenger comfort comparable to commercial aircraft, and be capable of all-weather operation. The plane should also provide for ease of maintenance and require only a single pilot. 
       SUMMARY OF THE INVENTION 
       [0005]    One embodiment consistent with the present invention includes a supplemental thrust system for an airplane comprising a submerged inlet duct, a diffuser section in communication with the inlet duct, and a duct outlet in communication with the diffuser section, an outlet nozzle in communication with the duct outlet, a turbocharger unit comprising one or more turbocharges in communication with each other to provide compressed air to the engine and the cabin, a heat exchanger unit comprising a plurality of heat exchangers in the diffuser section to provide cooling liquid to cool pressurized air in the cabin, engine intake air and the engine jacket where each turbocharger is coupled to the intake manifold of an engine to increase the power of the engine by introducing compressed air into the manifold. 
         [0006]    In another embodiment, the turbine exhaust of each of the turbochargers on each side converge into a single duct that is connected to the duct outlet. 
         [0007]    In another embodiment, an expansion duct is positioned between the duct outlet and the outlet nozzle. 
         [0008]    In another embodiment, the inlet duct is a NACA duct on the exterior of an aircraft. 
     
    
     
       DESCRIPTION OF THE DRAWINGS 
         [0009]    Details of the present invention, including non-limiting benefits and advantages, will become more readily apparent to those of ordinary skill in the relevant art after reviewing the following detailed description and accompanying drawings, wherein: 
           [0010]      FIG. 1A  depicts one embodiment of an aircraft consistent with the present invention; 
           [0011]      FIG. 1B  depicts a breakaway view of the aircraft of  FIG. 1 ; 
           [0012]      FIG. 1C  depicts a rear perspective view of the rear fuselage of  FIG. 1A ; 
           [0013]      FIG. 2  shows a top perspective view of the truss element; 
           [0014]      FIG. 3  depicts a breakaway view of the aircraft including the pressure vessel; 
           [0015]      FIG. 4  depicts one embodiment of one of the plurality of standoffs used to secure the pressure vessel; 
           [0016]      FIG. 5  depicts the attachment of skin to the truss elements; 
           [0017]      FIG. 6A  depicts the front landing gear affixed to the front bulkhead; 
           [0018]      FIGS. 6B-6E  depict the front landing gear retracting into the front fuselage; 
           [0019]      FIG. 7A  depicts the main landing gear connected to truss element; 
           [0020]      FIGS. 7B-7E  depict the main landing gear retracting into the rear fuselage; 
           [0021]      FIG. 8  depicts a heat recovery system used to increase the efficiency of the aircraft; 
           [0022]      FIG. 9  depicts a side view of the wing spar of the aircraft of  FIG. 1 ; 
           [0023]      FIG. 10A  depicts a flap control system included in the wing of the aircraft in  FIG. 1 ; 
           [0024]      FIG. 10B  depicts the flap control system with the plates removed; 
           [0025]      FIG. 11A  depicts the flap control system extending to lower the flaps; 
           [0026]      FIG. 11B  depicts the flap control system extending the flap downward; 
           [0027]      FIG. 11C  depicts the flap control system as it extends further outwards; 
           [0028]      FIG. 11D  depicts the flap control system with the foreflap and flap in the full extended position; 
           [0029]      FIG. 12  depicts the spoiler actuation system used to actuate the spoiler of  FIG. 9 ; 
           [0030]      FIG. 13A  depicts a trim actuator that is mechanically coupled to the elevator control system and similarly used in the dorsal fin control system; and 
           [0031]      FIG. 13B  depicts an interior view of the actuator along the lines A-A. 
       
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       [0032]    The purpose and advantages of the present invention will be set forth in and apparent from the description that follows, as well as will be learned by practice of the invention. Additional advantages of the invention will be realized and attained by the methods and systems particularly pointed out in the written description and claims hereof, as well as from the appended drawings. The term “top portion” is used herein to mean the portion of the fuselage farthest from the ground when the airplane is not in flight and the term “bottom portion” is herein to mean the portion of the fuselage closest to the ground when the airplane is not in flight. 
         [0033]    The design of the present invention makes use of aerodynamic shapes that are extensively laminar within their Reynolds number operating regime. Intersections of wing, empennage and fuselage are minimized, elliptical lift profiles are used on all lifting surfaces, and wing and horizontal tail shapes are approximately elliptical. The fuselage shape is derived from a modified zero camber extensively laminar airfoil section revolved about the longitudinal axis, thus making full use of pressure recovery to minimize form drag. The external aerodynamic shapes are mostly provided by gloves that fit over the frame of the aircraft, but are isolated from the frame so as to reduce surface waviness under load to an absolute minimum. This also permits easy one piece complete removal of the external skins for inspection of the frame and frame elements and maintenance of the operating systems attached to the frame. 
         [0034]    The wing structure of the aircraft consists of a box-and-channel structure that extends across approximately 90% of the span of the wing structure and is open to the rear but stabilized in compression. The wing structure is a composite beam with ply orientation and shape tailored to provide structural coupling in bending and torsion to generate variable wing washout as a function of bending to limit vertical wing loading and to provide damping of the major flutter modes. Tail surfaces have similar spar-and-glove design to allow for ease of inspection of all primary structure, decoupling of structural deformation from skin surfaces, and ease of exchange of external skin with new shapes for rapid repair of damaged surfaces as well as exchange of airfoil shapes with updated shapes or different internal systems should they become available. 
         [0035]      FIG. 1A  depicts one embodiment of an aircraft consistent with the present invention. The aircraft includes a forward fuselage  1 , a rear fuselage  2 , a midwing  3 , a vertical fin  4 , a ventral fin  5 , a horizontal tail  6  and a pusher type propeller  7 . The forward fuselage  1  and rear fuselage are covered in an external skin. The external skin may be made of a rigid fiber reinforced composite or metal such as, but not limited to, an aluminum alloy such as aluminum 2024 or aluminum 7078 or any other rigid material meeting a maximum waviness tolerance of 0.001 inches per inch measured over a two-inch span. 
         [0036]      FIG. 1B  depicts a breakaway view of the aircraft of  FIG. 1 . The frame of the aircraft includes a forward bulkhead  8  connected to an upper truss  9  on one end and two lower forward trusses  10  and  11  on an opposite side of the forward bulkhead  8 . The truss elements  9 ,  10  and  11  may be box type truss structures where the ends of the truss elements  9 ,  10  and  11  taper towards the forward bulkhead  8 , providing improved stiffness at the intersection of the truss elements  9 ,  10  and  11  and the forward bulkhead  8 . The truss elements  9 ,  10  and  11  are made of a rigid material including metal, fiberglass including S glass, or an equivalent material. Each composite truss element  9 ,  10  and  11  also includes a unidirectional upper cap, a unidirectional lower cap and shear/compression panels connecting the upper and lower caps. The shear panels may be comprised of +45/−45/0/90 plies of fiberglass, such as S glass or equivalent, configured for crush stiffness when loaded in vertical compression and for the minimal shear loading required by the triangulated configuration of the upper and lower caps. 
         [0037]    Each truss elements  9 ,  10  and  11  extends from the forward bulkhead  8  to the main bulkhead  15  where the truss elements  9 ,  10  and  11  are affixed to the main bulkhead  15  by fastener devices  12 ,  13  and  14 . The fastener devices  12 ,  13  and  14  may be comprised of transverse beams which may be formed of metal or a composite such as carbon fiber. Each fastener device  12 ,  13  and  14  is affixed to a respective truss element  9 ,  10  and  11  by a securing device such as a bolt passing through the fastening device  12 ,  13  or  14 , the corresponding truss element  9 ,  10  and  11  and a portion of the main bulkhead  15 . Each fastening device  12 ,  13  and  14  is attached to its respective truss element  9 ,  10  or  11  by wrapping the inner and outer plies of fastening device  12 ,  13  or  14  around the truss elements  9 ,  10  or  11  and doubling those plies back upon their outer and inner mating plies, respectively, thus mechanically locking the fastening device  12 ,  13  or  14  to respective truss element  9 ,  10  or  11 . Similar mechanical locking is used on the truss elements  19  and  20  of the rear fuselage. A main bulkhead transverse beam  16  is affixed to the exposed portions of the periphery of the main bulkhead  15  and is connected to the truss elements  9 ,  10  and  11 . 
         [0038]    Truss element  19  is affixed to the top portion of the main bulkhead  15  such that the central axis of the truss element  19  is substantially co-linear with the central axis of the truss element  9 . Truss element  20  is affixed to the bottom of the main bulkhead  15 , and truss elements  21  and  22  are affixed to opposing sides of the main bulkhead  15 . Each of the truss elements  19 ,  20 ,  21  and  22  may be box type beams. Truss elements  21  and  22  are configured to resist lateral loads induced by the vertical fin  5  and to provide support for skin cutouts required for the main landing gear doors and upper access hatches as described in further detail herein. 
         [0039]    Truss elements  19  and  20  extend from the main bulkhead  15  to a rear tail cone  24 . Each truss element  19  and  20  is affixed to the rear tail cone  24  using any known method of connection such as bolts, rivets or bonding. The upper surfaces, the surfaces facing away from the center portion of the aircraft, are coplanar with the surface of the tail cone  24 . The truss elements  21  and  22  are each affixed to a rear traverse bulkhead  25 , shown in  FIG. 1C , and to a forward traverse bulkhead  26 . A box section support  27 , shown in  FIG. 1C , is positioned between the rear traverse bulkhead  25  and forward traverse bulkhead  26  on the tail cone  24  to provide support for a vertical fin spar  28 . A horizontal tail spar  29  is positioned between the rear bulkhead  25  and an elevator bulkhead  34 , shown in  FIG. 1C . 
         [0040]    A fuel tank  33  is positioned adjacent the main bulkhead  15  in the rear fuselage  2 . The fuel tank  33  may be semicircular in shape and be positioned above the mid wing  3 . The fuel tank  33  is a separate replaceable bladder manufactured of a metal lined, highly damage tolerant composite structure that is internal to the fuselage and mounted on top of the wing spar, and is outside of the pressure vessel. Conventional wing tanks are difficult to seal and drain, and they are highly vulnerable to rupture in a crash due to their exposed distributed location along the wing span. With wing tanks, volumetric rearrangement in the event of crash-induced high G force loading is difficult to accomplish due to the walls of the tankage being part of the primary structure of the wing. By separably mounting the tank above the heaviest primary structure in the center of the aircraft, and by using a moderately volume-inefficient shape, volume rearrangement and thus survivability of the tank is enhanced. 
         [0041]      FIG. 1C  depicts a rear perspective view of the rear fuselage  2  of  FIG. 1A . The mid wing  3  is coupled to the main bulk head  15  by the sleeve  17 . The sleeve  17  is affixed to the main bulkhead  15  by a plurality of straps  39 . The straps  39  may be made of unidirectional fiberglass such as S glass, or any other material capable of securing the sleeve  17  to the bulkhead  15 . Each strap  39  extends around the periphery of the sleeve  17  such that a first portion of the strap  39  is in direct contact with the top surface of the sleeve  17 , a second portion of the strap  39  is in direct contact with a side surface of the sleeve  17 , and a third portion of the strap  39  is in direct contact with a lower portion of the sleeve  17 . A first end and second end of each strap  39  is affixed to the main bulkhead  15  by any known method of attaching a strap to a bulkhead, including rivets, bolts or bonding. 
         [0042]    A gusset  40  is attached to the lower portion of the sleeve  17  on one end and the main bulkhead  15  on the opposite end. The gusset  40  may be triangular in shape, with the wider portion of the gusset  40  connecting to the main bulkhead  15  and the narrower portion of the gusset  40  connecting to the bottom surface of the sleeve  17 . The gusset  40  acts to transfer upward loading force of the fuselage to the main bulkhead  15 . After installation, the fuel tank  33  shown in  FIG. 1B  may be positioned on the top surface of the sleeve or on a separate horizontal panel of transverse beam  37 , bonded to the structure. 
         [0043]    A transverse beam  37  is positioned on the bottom side of each truss element  21  and  22  and the side surface of the sleeve  17 . Half support ring  18  extends from the top surface of the transverse beam  37  adjacent to the truss element  21  to the top surface of the transverse beam  37  adjacent the lateral element  22 . The top surface of the half support ring  18  is substantially coplanar to the top surface of the truss elements  19 ,  20  and  21 . Full support ring  38  extends from one side of the truss element  20  to the opposite side of the truss element  20  such that the full support element connects to the truss elements  19 ,  21  and  22 . The top surface of the full support ring  38  is substantially coplanar with the top surfaces of the truss elements  19 ,  20  and  22 . Each support ring  18  and  38  is attached to truss elements  19 ,  20 ,  21  and  22  by multi-ply tabs as previously discussed or by any other method of attaching a support ring to a truss. Additional full and half support rings may be provided and affixed to the structure in a manner similar to the attachment of the half support ring  18  and full support ring  38 . 
         [0044]    The horizontal tail spar  29  is affixed between the rear bulkhead  25  and the elevator bulkhead  34 . The horizontal tail spar  29  is a continuous single piece spar that is pivotally attached to the rear fuselage by a pair of bearing units  36  mounted in a bearing carrier  35 . The outer sides of the bearing carrier  35  are affixed to the rear bulkhead  25  and the elevator bulkhead  34 . A tail wheel gusset  30  may be connected to the bottom surfaces of the rear bulkhead  25  and elevator bulkhead  34  to provide ventral fin and propeller protection from a tail strike due to over rotation during takeoff or landing. A wheel extension arm  31  and wheel  32  are rotatively affixed to one end of the gusset  30 . An actuator unit  33  is affixed to the bottom surface of the nose cone  24  between the gusset  30  and the end of the cone  24  such that the wheel extension arm  31  and wheel  32  can be extended during and retracted during flight. 
         [0045]      FIG. 2  shows a top perspective view of the truss element  20 . Truss element  20  includes a forward portion  552 , a rear portion  554 , support units  404  and a bulkhead connection plate  550 . The forward portion  552  and rear portion  554  are joined at center joint  560  and the support units  553  are affixed to the sides of the truss element  20  at the center joint  560 . The forward portion  552  and rear portion  554  are connected such that the top surface of the forward portion  552  and the top surface of the rear portion  554  form angle theta. In one embodiment, theta is approximately 180 degrees. In another embodiment, theta is between approximately 150 and approximately 178 degrees. 
         [0046]    The truss element  20  has a box structure with four sides and a hollow center portion. Openings may be cut along the sides of the truss element  20  to reduce the overall weight of the truss element  20  while also providing support for lateral and vertical loads encountered in flight, landing and takeoff conditions. The support units  553  extend from the sides of the truss element  20  at an angle beta relative to the top surface of the truss element  20 . Each support unit  553  includes a connection plate  410  on the end of the support unit  404  furthest from the truss element  20 . The bulkhead connection plate  550  is affixed to the front surface of the truss  20 . The bulkhead connection plate  550  includes a substantially arc shaped portion that is shaped to engage a lower portion of the main bulkhead  15  using connection openings  551 . A plurality of sidewall connection openings  555  are positioned along the sidewalls of the truss element  20  for connecting a motor mount to the truss element  20 . 
         [0047]      FIG. 3  depicts a breakaway view of the aircraft including the pressure vessel. The pressure vessel  43  is positioned in the forward fuselage assembly  42  between the main bulkhead  15  and the nose of the aircraft. Because of the differing forms of the loads induced by local loading by payloads, aerodynamic loads and ground loads and the distributed loading from pressurization, payload-induced loading is applied to the fuselage truss elements  9 ,  10 ,  11 ,  19 ,  20 ,  21  and  22 , and not the pressure vessel  43 , which is isolated from the truss elements  10 ,  11 ,  19 ,  20 ,  21  and  22 . Isolating the pressure vessel  43  eliminates waviness of the external skin due to pressure deflections as would be the case with a conventional monocoque aircraft fuselage structure. Minimal waviness is a necessary criterion for the maintenance of laminar flow over the fuselage, resulting in corresponding low parasite drag of the fuselage. 
         [0048]    The pressure vessel  43  is positioned in the forward fuselage assembly  42  such that it is surrounded by the truss elements  9 ,  10 , and  11  and main bulkhead  15 . The pressure vessel  43  is structurally isolated from the truss by padding rings on the truss elements  9 ,  10  and  11  that support the pressure vessel  43 . Vertical deflection of the truss elements  9 ,  10  and  11  will not couple to the pressure vessel  43 , and as a consequence structural loading of the elements  9 ,  10  and  11  by payloads will produce essentially no induced loads in the pressure vessel  43 . Similarly, pressurization of the pressure vessel  43  will contribute no loading to the truss elements  9 ,  10  and  11  in any direction because the two structures are completely decoupled via the pads. The pressure vessel  43  is indexed to the truss elements  9 ,  10  and  11  by a single standoff (not shown) that penetrates the pressure vessel  43  through a close tolerance hole and is sealed to internal pressure of the pressure vessel  43  by a circular seal that is free to slide in the radial direction on the standoff. The indexing standoff (not show) is one of a number of standoffs that penetrate the pressure vessel  43  through oversized reinforced holes in the pressure vessel  43  and which carry the loads sustained by the floorboards, internal panels and other internal appurtenances through the pressure vessel  43  outwards into the truss elements  9 ,  10  and  11 . All but two of these reinforced holes are loose longitudinal and circumferential fits to the standoffs to allow for pressure vessel expansion, and thus there is only a single longitudinal and circumferential locating position. 
         [0049]    The parts of the pressure vessel  43  forward and aft of an index position are free to expand and contract longitudinally, circumferentially and radially without coupling any loads or deflections into the truss elements  9 ,  10  and  11  and conversely, truss element deflections cannot produce induced loading in the pressure vessel  43 . The front dome of the pressure vessel  43  is an ideal hemispherical shape with cutouts for a windshield and windows. Those cutouts are ring and strap reinforced to resist the tangential pressure loads, and the panes are coupled to the vessel  43  in only a radial direction. Therefore, no circumferential loads are transmitted. 
         [0050]    The differential thermal expansion and the pressure-induced diaphragm deflections of the panes from the pressure vessel  43  are also reduced by the ring and strap reinforcement. In contrast, the doors are set coplanar to the pressure vessel  43  walls and are fastened in a tangentially load bearing semi-continuous fashion to the walls of the pressure vessel  43  around their entire circumference by means of the sealing device  67 . Internal pressure increases latching forces of the doors to the walls of the pressure vessel  43 . The doors are thus load-bearing elements of the pressure vessel  43 . 
         [0051]      FIG. 4  depicts one embodiment of one of the plurality of standoffs used to secure the pressure vessel. The standoff includes two load distribution plates  63  and  69 . The external plate  69  is affixed to a truss element  9 ,  10  or  11 . The interior plate  63  is affixed to a load bearing structure within the pressure vessel  43 . A cylindrical standoff  64  has opposing ends fastened to the distribution plates  63  and  69  by fasteners  70 . The fasteners  70  are configured to carry the full load applied to the standoff, and are held in positioned by a locking mechanism such as a tabbed washers, safety wire or any other means of locking the fasteners  70  in place. The cylindrical standoff  64  extends through an opening in the wall of the pressure vessel  62 . The opening in the wall of the pressure vessel  43  is sized to accommodate the expansion and contraction of the pressure vessel  43 , and the movement of the pressure vessel  43  during operation of the aircraft. Two standoffs  64  that are diametrically opposed, are connected to openings in the pressure vessel  43  that do not compensate for expansion and contraction of the pressure vessel  43  during operation. 
         [0052]    The openings in the pressure vessel  43  are reinforced by a plate  65  that has a surface coplanar to the outer surface of the pressure vessel  43 . The plate  65  may be made of any material capable of withstanding tangential loads of the pressure vessel  43  including steel, aluminum and alloys thereof, carbon fiber or any other material that can withstand the tangential loads of the pressure vessel  43 . The material of the plate  65  also has thermal expansion and elastic characteristics comparable to the material used in the pressure vessel  43 . In one embodiment, the pressure vessel  43  and the plate  65  are made from the same material. The interior portion of the plate  65  engages a washer  66 . The washer  66  includes a cylindrical boss sized to accommodate a sealing device  67 , such as an O-Ring. The sealing device  67  engages the cylindrical standoff  64  such that the washer  66  is in direct contact with the cylindrical standoff  64 . A spring  68  positioned between the plate  69  and the washer  66  forces the washer  66  against the plate  65 . 
         [0053]    The cylindrical standoffs  64  penetrate the pressure vessel  43  through the openings in the pressure vessel  43  wall which reinforced by the washer  66 -spring  68  combination to carry the tangential pressure induced loads. The standoffs  64  are fastened to truss elements  9 ,  10  and  11  as necessary for load distribution. The standoffs  64  are pressure sealed to the wall of the pressure vessel  43  by means of the washers  66  and spring  67 , which bosses are sealed by the sealing device  67  that seals the washers  66  to the cylindrical standoffs  64  by the washers&#39;  66  flat but flexible surface resting on the corresponding flat surfaces provided on the inside of the wall of the pressure vessel  43 . The combination washer  66  and spring  67  are free to slide both on the standoff  64  outer diameter and on the flat on the inside of the pressure vessel  43  wall. The internal diameter of each opening is large enough with respect to the outer diameter of the penetrating standoffs  64  to allow for all anticipated expansion and contraction of the pressure vessel  43  and deflections of the truss under load. Using these techniques, the pressure vessel  43  sees only well distributed loading due to internal pressure and is completely isolated from payload-induced loads and other flight and ground loads. The weight of the pressure vessel  43  itself is supported by elastomeric foam attached to the interior surfaces of the beams of the forward truss elements  9 ,  10  and  11 . This provides only a padded resting surface for the exterior of the wall of the pressure vessel  43 . The pressure vessel  43  can be installed and removed from the forward fuselage  41  as a unit. This is done by separating the forward  41  and rear  42  halves of the fuselage and inserting or removing the pressure vessel through the rear opening of the forward fuselage. 
         [0054]    The internal dimensions of the forward fuselage truss elements  9 ,  10  and  11  are slightly larger than the maximum pressurized diameter of the pressure vessel  43 . The truss elements  9 ,  10  and  11  are bonded to the exterior skin of the aircraft, and the skin forms a shear web between the top truss element  9  and the bottom truss elements  10  and  11 . The truss elements  9 ,  10 , and  11  are bonded to the forward bulkhead  8  in a triangulated fashion, and the forward bulkhead carries the nose gear loads into the truss elements  9 ,  10  and  11 . By using multiple standoff penetrators to carry the loads from inside the pressure vessel  43 , to the truss elements  9 ,  10  and  11 , a relatively uniformly distributed load on the truss elements  9 ,  10  and  11  is achieved. This minimizes local deflections and high stress points that could induce undesirable waviness into the outer skin of the fuselage. Both the floorboard structure and the box beams that form the bottom elements of the truss are used as crush structure to manage energy absorption to enhance crashworthiness. The overall aircraft structure is designed for  26  g ultimate loading. 
         [0055]    The external skin of the forward fuselage is composed of a formed sandwich panel which is bonded to the truss elements  9 ,  10 , and  11 , the forward bulkhead  8  and an attachment ring at the rear of the forward fuselage. The rear fuselage skin is similar and is bonded to the upper, lower, and side truss elements  19  and  20 . The rear half of the fuselage contains the main bulkhead  15 , which is bonded to the forward ends of the truss elements  19 ,  20 ,  21  and  22  and the rear skin. The sleeve  17  is bonded to the main bulkhead  15  and to two truss elements  21  and  22  which are likewise bonded to the skin and to the main bulkhead  15 . The truss elements  21  and  22  are provided to stiffen the rear fuselage in the lateral direction. This is necessary due to the large skin cutouts for the main landing gear doors and other access hatches. 
         [0056]    The truss elements  19  and  20  are single box beams on both top and bottom. All four box beams and the rear fuselage  43  skin are bonded to the tail cone  24  which carries the horizontal and vertical tail surface attachments and bearings. To allow for a sliding seal surface between the two halves of the horizontal tail and the fuselage, the tail cone  24  is surrounded by a removable, mechanically-fastened fairing that is contoured to fit the rotational movement of the inner surfaces of the horizontal tail. This fairing is a replaceable wear surface that provides the sealing surface for the sliding seal between the horizontal tail and the fuselage. 
         [0057]      FIG. 5  depicts the attachment of skin to the truss elements. The forward skin  44  is bonded to a box ring  49  with a core  50 . The rear skin  45  is bonded to the main bulkhead  15 , the main bulkhead  15  includes a forward skin  47 , a rear skin  48 , and a core  46 . Doubler plies or metal doublers  51  and  52  provide stress distribution of the local loading generated by the fasteners,  55 . There are a multiplicity of fasteners distributed circumferentially around the box ring  49  to provide a semi-continuous engagement between the forward skin  44  and the rear skin  45 . The fasteners  55  are shoulder bolts that provide shear coupling between the skins, as well as adequate tensile coupling. The fasteners  55  are threaded into a sealed nut plate  53  with a counter bored section to engage the shoulder of the fastener  55 . To prevent crushing of the core of the main bulkhead  15 , a tubular standoff is bonded to the forward skin of the bulkhead,  47  and the rear skin of the bulkhead  48 . This allows the fastener  55  to load the forward bulkhead skin  47  against the rear doubler  52  the rear fuselage skin plies  45  the box ring plies  49  the forward fuselage skin plies  44  and the forward doubler  51  stacked in that order without crushing the main bulkhead core  46  or the box ring core  50 . 
         [0058]      FIG. 6A  depicts the front landing gear  43  affixed to the front bulkhead  8 . The landing gear  43  may be an oleo type trailing link landing gear.  FIG. 6B-6E  depict the front landing gear retracting into the front fuselage.  FIG. 6B  shows the landing gear  43  in the fully extended position. The front landing gear  43  includes an actuation device  612 , a wheel  602 , a swing arm  604 , a forward link arm  606 , a horizontal hinged plate  608  and an oleopneumatic cylinder  610 . The swing arm  604  includes two parallel plates with one end of each plate being connected to the wheel  602  by an axle that passes through the center of the wheel  602  and through corresponding openings in the plates of the swing arm  604 . The other end of the swing arm  604  opposite the wheel  602  is rotatively coupled to the forward link arm  606  by a pin  610  that allows the swing arm  604  to rotate relative to the forward link arm  606 . 
         [0059]    The hinged plate  608  is rotatively coupled to the bulkhead  8  by hinges  612  connected to the bulkhead  8  such that the plate  608  is pulled towards the bulkhead  8  as the landing gear  43  is moved to the refracted position and the plate  608  is moved to a position substantially perpendicular to the bulkhead  8  when the landing gear  43  is fully extended. The oleopneumatic cylinder  610  may be a hydraulic piston or air filled piston. The oleopneumatic cylinder  610  has a first end connected to the swing arm  604  between the wheel  602  and the forward link arm  606 . In one embodiment, the oleopneumatic cylinder  610  is connected at approximately the center of the swing arm  604 . The oleopneumatic cylinder  610  passes through the plate  608  allowing the second end of the oleopneumatic cylinder  610  to rotatively connect to the bulkhead  8  such that the oleopneumatic cylinder  610  rotates towards the bulkhead  8  as the landing gear  43  is retracted. The forward link arm  606  is rotatively connected to the oleopneumatic cylinder  610  at a positioned just below the plate  608 . The actuation device  612  is rotatively coupled to the bulkhead  8  by a hinge and to the plate  608  by a hinge. The actuation device  612  includes a base portion  614 . The actuation device  602  may be a hydraulic actuator, a linear actuator or any other device capable of retracting and extending the landing gear  43 . 
         [0060]      FIG. 6C  depicts the landing gear  43  as the landing gear  43  is retracted into the fuselage. The actuation device  612  is activated such that the extension arm  614  retracts into the actuation device  612  pulling the plate  608  towards the bulkhead  8 . As the plate rotates towards the bulkhead  8 , the forward link arm  606  rotates towards the plate  608  and the swing arm  604  rotates towards the forward link arm  606  pulling the wheel  602  upward.  FIG. 6D  depicts the landing gear  43  retracting into the fuselage. As the actuation device  612  continues to pull the extension arm  614  into the base  600 , the plate  608  is pulled further towards the bulkhead  8  causing the oleopneumatic cylinder  610  to rotate upward and compress, and the extension arms  604  and  606  rotates towards the plate  608  pulling the wheel  602  upwards into the fuselage.  FIG. 6E  depicts the landing gear  43  fully retracted into the fuselage. The landing gear is extended by extending the extension arm  614  out of the actuation device  612  such that the plate  608  rotates away from the bulkhead  8 . 
         [0061]      FIG. 7A  depicts the rear landing gear  700  connected to truss element  20 . The rear landing gear  700  includes two frames  702  that are each substantially A-shaped. Each frame  716  is rotatively affixed to a side of the truss element  20  by a pin. Each frame  716  is also rotatively connected to a trailing link  704  by a pivot joint  706 . The pivot joint  706  is substantially ‘C’ shaped and is sized to accommodate an end of the trailing link  704 . A pin  708  passes through both sides of the pivot joint  706  and the trailing link  704  to secure the trailing link  704  in the pivot joint  706 . The opposite end of the trailing link  704  is connected to a wheel  710  and one end of a cylinder  712 . The other end of each cylinder  712  is rotatively connected to a support unit  553  on the truss element  20  via a universal joint. 
         [0062]    Each frame  702  includes an overcenter locking unit  714  that is configured to secure the frame in a fully extended position and a support plate  716  rotatively connected to the truss element  20  by a hinge. The end of the locking unit  714  furthest from the truss element  20  is rotatively coupled to the end of the support plate  716  furthest from the truss element  20 . Each locking unit  714  is separated into two sections by a pin. The cylinder  712  may be an hydraulic piston filled with a hydraulic fluid and air. The cylinder  712  includes a cylinder body  718  and rod  720  extending from the cylinder body  718 . 
         [0063]      FIG. 7B  depicts the rear landing gear  700  in the fully extended position. The locking units  714  are fully extended such that the support plate  716  is substantially perpendicular to the side of the truss element  20 .  FIG. 7C  depicts the rear landing gear  700  retracting into the fuselage. A retraction cylinder folds the locking units  714 , pulling the support plate  716  upward. As the support plate  716  moves upward, the two portions of the locking unit  714  rotate about the pin, separating the two portions of the locking unit  714  such that the two portions of the locking unit  714  move towards each other. The movement of the cylinder  712  causes the trailing link  704  to rotate towards the truss element  20 , bringing the wheels  710  towards the fuselage.  FIG. 7D  depicts the rear landing gear  700  further retracting into the fuselage. As the support plate  716  continues to move towards the truss element  20 , the cylinder pulls the wheels  710  into the fuselage. 
         [0064]      FIG. 7E  depicts the rear landing gear  700  fully refracted into the fuselage. The rod  720  is fully extended out of cylinder  712 , and the support plate  716  and the central axis of the wheel  710  both are substantially parallel to the side of the truss element  20 . The two portions of the locking unit  714  are separated by an angle with the angle being less than 90 degrees. 
         [0065]    Propulsion of the aircraft may be provided by a fixed-pitch eight blade composite blade propeller mounted at the rear of the fuselage on the centerline axis. The propeller airfoil sections and section incidence angles are configured to provide maximum efficiency at cruise at 50,000 ft. altitude and above. Propeller diameter is also optimized for the high altitude cruise environment and as a result essentially eliminates supersonic blade velocities during low altitude operation. The optimum propeller diameter is slightly smaller than maximum fuselage diameter which coincidentally reduces the probability of bird strike and other foreign object damage. 
         [0066]    The propeller is connected to two engines by a drive shaft extending from the output shaft of a gear box. The engines are liquid-cooled diesel engines driving torque converters connected to the gear box. Multi-stage turbo charging is provided to compensate for altitude and to provide cabin pressurization. Engine heat exchangers, turbo chargers and intercooler heat exchangers are all mounted in ducts configured to provide thermal recovery of waste heat for supplemental propulsion. Engine exhaust is likewise used in the rear of the same duct to provide an injection pump function both for cooling air circulation during low speed operation and to provide additional thrust during flight. 
         [0067]    The torque converters are provided to isolate the propeller, drive shaft, and gear box from periodic variations of engine torque and to provide for necessary torque multiplication required by the propeller during low speed operations. Traditional propeller and engine combinations provide no vibration isolation and match engine torque output to propeller demands by varying the pitch of the propeller to reduce the propeller torque demand. This results in much higher propeller speeds during near ground operations, and consequently much greater noise output, and it also results in a propeller airfoil and pitch distribution that is never optimum. The use of torque converters without lockup clutches allows an engine shutdown to disconnect the inoperative engine from the driveshaft and propeller. In the event that both engines are shut down, the propeller is completely disconnected from both engines. Alternators and emergency cabin pressurization remain connected to the drive shaft and are driven by the wind milling propeller. This is the only external mechanical drag load applied to the propeller aside from bearing friction and freewheeling transmission friction. 
         [0068]      FIG. 8  depicts a heat recovery system  800  used to increase the efficiency of the aircraft. Cooling air is introduced to the heat recovery system  800  from ducts  802  located on the exterior of the aircraft. The ducts may be NACA submerged ducts. The air introduced via the ducts  802  passes over a first heat exchanger  804 . The first heat exchanger  804  provides cool fluid used to cool the air bled from the turbo charger used to pressurize the cabin. The air then passes over a second heat exchanger  806  that provides cooling liquid for the intercoolers that cool the engine air intake. The air then passes through a third heat exchanger  808  that cools the liquid from the engine jacket. 
         [0069]    After leaving the third heat exchanger  808 , the air passes across the turbo chargers  810 . The output of the turbo chargers  810  are connected to the manifold  812  and intercoolers of the engine to provide compressed air to the engine to increase the thrust produced by the engine. The turbine exhaust of all turbo chargers on each side are combined into a single tubular exhaust pipe  814  which combines with a convergent part of the duct  816  to form an injection pump that mixes the turbine exhaust with the heated cooling air flow and then flows through a nozzle to provide additional thrust. In one embodiment, the thermal recovery system  800  generates an additional 5-6 pounds of thrust. 
         [0070]      FIG. 9  depicts a side view of the wing spar  900  of the aircraft of  FIG. 1 . The wing skin  902  and a sleeve  904  are bonded to the skin  902  at upper and lower surfaces and at corners  906  of the sleeve  904 . The sleeve  904  is a tight fit to the wing spar  900  and is pinned to the spar  900  at the wing root by a pin located on the neutral axis of the spar  900 . A spoiler  910  and vent  912  are provided for roll control and flight path control. The spoiler  910  and vent  912  are linked to open together to provide a slot lip type aileron. The wing skin  902  is bonded internally to the sleeve  904  such that the skin  904  that slips over the outside of the spar  900  to form a close fit to the spar  900  that is free to slide in the span wise direction to accommodate flexure of the spar  900 . In one embodiment, the skin  902  is fastened to the spar  900  at the wing root only. By securing the skin  902  to the spar  900  at the wing root only, the skin  902  is isolated from the spar  900  in order to minimize skin  902  buckling due to bending and to allow for quick replacement of damaged skin sections  902 , ease of updating of wing systems and airfoil shapes, and quick installation and removal for inspection of the spar  900  structure and the flap and spoiler systems. 
         [0071]      FIG. 10A  depicts a flap control system  1000  included in the wing of the aircraft in  FIG. 1 . The flap control system  1000  includes a plurality of control stations  1001  that each includes a plurality of plates  1002 ,  1004 ,  1006  and  1008  connected together by fasteners  1010  passing through the corners of each plate. Each plate  1002 ,  1004 ,  1006  and  1008  includes an opening  1012  that is sized to accommodate a drive shaft  1014 . Each station  1001  is secured to the wing spar  900 . The drive shaft  1014  extends the length of the wing and is connected to each control station  1001 . The plates  1004  and  1006  have a length longer than the plates  1002  and  1008 . One end of the plates  1004  and  1006  includes an opening  1016  that is sized and shaped to accommodate a fore flap  1018 . The fore flap  1018  is connected to a flap  1020  by a flap plate (not shown). 
         [0072]      FIG. 10B  depicts the flap control system  1000  with plates  1002 ,  1004  and  1008  removed. A chain  1050  is driven by the drive shaft  1014  connected to a sprocket  1052 , and which wraps around idler gears  1054  and  1056 . The drive shaft  1014  rotates both clockwise and counterclockwise to drive the chain  1050  in both forward and reverse directions to extend and retract the flap  1020 . The chain  1050  is tensioned by the idler gears  1054  and  1056  and is attached to chain shoe  1058 . The chain shoe  1058  is positioned and slides in slot  1060  on the inner surface of plate  1008  and is rotatively connected to one end of a support arm  1062  such that the chain shoe  1058  can rotate relative to the support arm  1062 . The opposite end of the support arm  1062  connects to the foreflap  1018  through a slot  1064 . A second shoe  1068  is connected to the support arm  1062  at approximately the center of the support arm  1062 . The second shoe  1068  is positioned and slides in slot  1070  in plate  1008 . Slot  1070  is substantially arc shaped and is positioned to allow optimum positioning of the flap  1020  or foreflap  1018  with respect to the wing. A link arm  1072  is substantially ‘U’ shaped and is connected to the second shoe  1068  at substantially the center of the link arm  1072 . One end of the link arm  1072  is coupled to a third shoe  1074  that is positioned and slides in a slot  1076 . Slot  1076  is substantially arc shaped and is positioned below the slot  1070 . The end of the link arm  1072  opposite the end connected to the third shoe  1074  is connected to tilt arm  1078 . The end of the tilt arm  1078  not connected to the link arm  1072  is connected to the lower portion of the flap plate  1066  at a position below the connection of the support arm  1068  to the flap plate  1066 . 
         [0073]      FIG. 11A  depicts the flap control system  1000  in the retracted or zero degree position. The chain shoe  1058  is positioned adjacent to the idler gear  1056  in the slot  1060 , the third shoe  1074  is positioned near the bottom edge of the plate  1008  in the slot  1076  and the tilt arm  1078  is in its full refracted position.  FIG. 11B  depicts the flap control system  1000  extending the flap  1020  downward. The sprocket  1052  drives the chain  1060  moving the sprocket  1052  towards the flap  1020 . As the sprocket  1058  moves, the support arm  1062  pushes the foreflap  1018  and the flap  1020  outwards. As the support arm  1062  moves, the link arm  1072  moves in the slot  1076  pulling the tilt arm  1078  inwards causing the flap plate  1066  to rotate in a clockwise manner. 
         [0074]      FIG. 11C  depicts the flap control system  1000  as it extends further outwards. As the chain  1050  continues to move the chain shoe  1058  the support arm  1062  pushes and rotates the foreflap  1018  and the link arm  1072  continues to move in the slot  1076  to push the tilt arm  1078  away from the plate  1008  to rotate the foreflap  1018  and flap  1020  down.  FIG. 11D  depicts the flap control system with the foreflap  1018  and flap  1020  in the full extended position. The chain shoe  1058  is positioned in the portion of the slot  1060  furthest outward. The link arm  1072  is positioned in the slot  1076  such that a portion of the link arm  1072  is substantially perpendicular to the tilt arm  1078 . The flap  1020  is positioned such that the training edge of the flap  1020  points substantially downward. 
         [0075]    The flap control system may be a 90% span double-slotted flap system including slot lip spoilers and spoiler vents used for roll control and glide path modulation. All flap tracks are fully internal to the wing when the flaps are refracted, and extension is by means of drive shaft  1014  extending across the full 90% of span with the drive shaft actuator in the center of the wing. Each control station  1001  along the wing converts rotational motion of the drive shaft  1014  to linear motion of the support arm  1062  and the link arm  1072  and the motion of the tilt arm  1078  by means of the sprocket  1056  and chain  1050 . The tooth count of each sprocket  1056  is a fixed ratio to chord length of the wing at each span wise station. 
         [0076]      FIG. 12  depicts the spoiler actuation system  1200  used to actuate the spoiler  91  of  FIG. 9 . The spoiler  910  is actuated by means of two slotted mount plates  1202  and  1204  plates and a cam plate  1205  to provide positive control of extension and retraction of the spoiler  91  and full lock of the spoiler  910  in the refracted position. Normally, the cam plates  1205  are linked together and move synchronously, locking one spoiler in the locked down position while proportionately deploying the opposite spoiler with respect to the yoke rotation. Approach path modulation is provided by moving the cam plates  1205  on opposite wings either closer together or farther apart with respect to one another. The entire flap and spoiler mechanism is mounted in the open rear half of the spar of the wing, which provides unrestricted access to the mechanism when the wing glove is removed. 
         [0077]      FIG. 13A  depicts a trim actuator  1300  that is mechanically coupled to the elevator control system. A similar actuator is used on the dorsal fin control system. The actuator includes a base housing  1302  and an extension rod  1304  that slides into and out of the base housing  1302 . The end of the extension rod  1304  opposite the base housing  1302  and the end of the base housing  1302  opposite the extension rod  1304  each includes a securing unit  1306  and  1308  affixed to the end thereon. The securing units  1306  and  1308  may be eyelets. 
         [0078]      FIG. 13B  depicts an interior view of the actuator  1300  in a centered, compressed and extended position. The base housing  1302  contains two springs  1310  and  1312  and a stop  1314  fastened to the cylinder bore. The extension rod piston  1304  engages two washers  1316  and  1318  that lie on either side of the stop  1314  and against which the springs  1310  and  1312  rest. When the extension piston  1304  is moved in either direction from its neutral position aligned with the stop  1314 , it compresses one of the springs  1310  and  1312  which drives the extension rod piston  1304  back into the neutral position. The overall position of the actuator is controlled by a ball bearing jack screw that sets the trim position of the elevator, and a second similar system sets the position of the dorsal fin. The surfaces of the extension rod  1304  and base housing  1302  are never in a stick-free condition, thus eliminating the need for geared tabs and other complications for stabilization. 
         [0079]    The aircraft cabin may be approximately 74 inches high and include an approximately 78 inch width having a minimum 50 inch seat pitch. The aircraft has a service ceiling of approximately 65,000 feet, and a normal cruise speed of between approximately 460 to approximately 510 mph, with a specific fuel consumption of approximately 30 to approximately 42 mpg depending on cruise speed and altitude. Landing stall speed is approximately 70 mph, takeoff and landing speeds are approximately 90 mph, and runway requirements are approximately 3000 ft. 
         [0080]    It is to be understood that both the foregoing general description and the following detailed description are exemplary and are intended to provide further explanation of the invention claimed. The disclosed configuration is the preferred embodiment and is not intended to preclude functional equivalents to the various elements. 
         [0081]    The accompanying drawings, which are incorporated in and constitute part of this specification, are included to illustrate and provide a further understanding of the invention. Together with the description, the drawings serve to explain the principles of the invention.