Patent Publication Number: US-2012037751-A1

Title: Supersonic flying wing

Description:
FIELD OF THE INVENTION 
     The present invention relates to a method and system for a multidirectional flying wing having a reduced sonic boom and improved subsonic and supersonic aerodynamics efficiency. 
     BACKGROUND OF THE INVENTION 
     Supersonic commercial flight has always been a great interest of aircraft design engineers, scientists, and business professionals due to the potential to reduce inter-continental travel time. Currently, a trip from Chicago to Hong Kong on commercial flights today takes approximately 14 hours. A supersonic transport (SST) may be able to make a similar journey in about 5 hours. This significantly shorter flight time makes SSTs very attractive. However, building a practical SST presents several challenges. 
     First, an SST generates a sonic boom. As it travels through the air, an SST generates a shock wave in the air. This shock wave expands out as a pressure wave that even at great distances may be as strong as 3 pounds per square foot (psf.) or more. The resulting sonic boom is very loud, and may even be strong enough to break glass or other brittle materials. The Aérospatiale-BAC Concorde®, the only SST ever to be used commercially use, was forbidden to fly at supersonic speeds over land due to the undesirable and problematic sonic boom it generated. Therefore, a practical SST must have a weak sonic boom. 
     Second, all supersonic aircraft experience significant wave drag caused by air traveling at supersonic speeds over the aircraft. Wave drag may be four or more times as strong as drag experienced at subsonic speeds. Wave drag significantly decreases the efficiency of the aircraft and vastly increases fuel consumption, strain on the aircraft and operating costs. Therefore, a practical SST must have a reduced wave drag. 
     Third, aircraft designed to fly at supersonic speeds typically perform poorly at slower, subsonic speeds. At subsonic flying speeds, such as during take-off and landing, aircraft having a high aspect ratio, low sweep angle and blunt leading edge on the airfoil perform better. The aspect ratio is the ratio of the wingspan to the wing chord, i.e. the width of the wing. The wing sweep angle is the angle at which the wings deviate from perpendicular to the direction of flight. The leading edge is the front of the wing. Aircraft designed for subsonic flight typically have a relatively narrow wings protruding directly at or close to perpendicular from the direction of flight and have a blunt leading edge. 
     Aircraft at supersonic speeds perform better with a lower aspect ratio (shorter wingspan), a relatively high sweep angle and a sharp leading edge on the airfoil. As a result, supersonic aircraft tend to perform poorly at subsonic speeds and are more difficult to control during take-off and landing. 
     Several attempts have been made to overcome these difficulties. For example, the Grumman F-14 Tomcat® incorporates wings having a variable sweep angle, allowing it to adjust the craft&#39;s aspect ratio and wing sweep. However, variable sweep wings require the inclusion of the sweep angle varying mechanisms, substantially increasing the weight of the plane. 
     Others have modified the nose of a plane to reduce the sonic boom. Specifically, the nose is made very blunt and rounded to distribute the shock wave. Modification of the shape of the lower portion of the front of the fuselage decreases the sonic boom experienced on the ground. However, the reduction in sonic boom is insufficient to allow safe use of supersonic aircraft over populated areas. 
     The Gulfstream Quiet Spike® has also been developed to decrease the sonic boom. This concept employs a spike added to the front of a fuselage. The spike increases in diameter in a stepwise manner, thereby creating several weak shock waves. This mechanism creates more, smaller shock waves. However, the spike usually extends at least a third of the length of the aircraft, making it unstable and impractical, and does not reduce the sonic boom to an acceptably low level. 
     For years NASA® and others experimented with an oblique wing design. The oblique wing consisted of a wing mounted to the fuselage in a pivotable fashion. As the plane increased in speed, the wing pivoted. This design provides decreased wave drag and improved aerodynamic performance. However, the asymmetric wing sweep, forward on one side of the plane and backward on the other, couples pitch with roll and decreases the stability of the craft. Naturally occurring flight designs, such as birds, insects and even dinosaurs, have symmetric planforms (the shape of a flying craft as viewed from directly above). Thus, millions of years of evolution also argue against asymmetric planforms. 
     It is therefore desirable to provide a supersonic transport aircraft. It is also desirable to provide a supersonic transport aircraft having reduced sonic boom, wave drag, and that operates efficiently at subsonic speeds. 
     SUMMARY OF THE INVENTION 
     The present invention advantageously provides an aircraft comprising a lifting body and a propulsion element. The lifting body will rotate about the propulsion system. The propulsion system will stay in the flight direction to provide the required thrust and not rotate. 
     In another aspect the invention provides a supersonic bidirectional flying wing comprising a lifting body having bilateral symmetry across a longitudinal plane and bilateral symmetry across a transverse plane, a substantially isentropic compression bottom surface, an upper surface, a propulsion element rotatably attached to the upper surface, one or more engines mounted on the propulsion element, a front longitudinal tip and a back longitudinal tip, each pivotably coupled to the lifting body such that they may pivot to approximately perpendicular to the isentropic compression bottom surface during flight in the transverse direction, wherein the lifting body has a sweep angle of at least 40 degrees and an aspect ratio of less than 5 during flight parallel to a longitudinal axis and wherein the lifting body has a sweep angle less than 50 degrees and an aspect ratio of at least 3 during flight parallel to a transverse axis. 
     In another aspect the invention provides a method for subsonic and supersonic flight with reduced sonic boom generation comprising providing a lifting body having a substantially isentropic compression bottom surface, wherein the lifting body has a high aspect ratio and low sweep angle in a transverse direction and has a low aspect ratio and a high sweep angle in a longitudinal direction, propelling the lifting body in the transverse direction during subsonic flight and propelling the lifting body in the longitudinal direction during supersonic flight. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       A more complete understanding of the present invention, and the attendant advantages and features thereof, will be more readily understood by reference to the following detailed description when considered in conjunction with the accompanying drawings wherein: 
         FIG. 1  is a perspective view of a supersonic bidirectional flying wing constructed in accordance with the principles of the present invention; 
         FIG. 2  is the planform of the supersonic bidirectional flying wing shown in  FIG. 1 ; 
         FIG. 3  is an alternative planform of a supersonic bidirectional flying wing constructed in accordance with the principles of the present invention; 
         FIG. 4  is a perspective view of an alternative supersonic bidirectional flying wing constructed in accordance with the principles of the present invention; 
         FIG. 5  is the longitudinal cross-section of the supersonic airfoil of the supersonic bidirectional flying wing shown in  FIG. 1 ; 
         FIG. 6  is the transverse cross-section of the subsonic airfoil of the supersonic bidirectional flying wing shown in  FIG. 1 ; 
         FIG. 7  is a diagram of a slat and airfoil used in accordance with the principles of the present invention; 
         FIG. 8  is a diagram of a radial airflow ejector and airfoil used in accordance with the principles of the present invention; 
         FIG. 9  is a diagram of a radial airflow and a unidirectional airflow; 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     The present invention provides for a supersonic aircraft that is capable of flying at subsonic speeds in one direction and then rotating 90° to fly at supersonic speeds in another direction, thereby reducing wave drag and increasing efficiency. The invention includes a substantially isentropic compression lower surface that results in a weak or no downward shock wave and a reduced sonic boom. 
     The aircraft rotation will be controlled by split flaps (similar to that of a B2 bomber) on the wings, which will split and open on one side of the flying wing and create the yaw moment. There will be a lock/unlock mechanism between the flying body and the propulsion system. When it is unlocked between the flying body and the engines, the yaw moment will rotate the flying body and the propulsion system will remain in the flight direction. When the flying body and the engines are locked, the yaw moment will rotate the whole system as the yaw control. The pitch will be controlled by the flaps or the ailerons on the two sides of the flying axis with both sides deflecting in the same direction. The rolling will be also controlled by the flaps or ailerons with both sides deflecting in the opposite direction. 
     Referring now to the drawing figures in which like reference designators refer to like elements, there is shown in  FIG. 1  an embodiment of a supersonic bidirectional flying wing constructed in accordance with principles of the present invention and generally designated as “ 10 .” The flying wing  10  is shown in its subsonic configuration, flying in a transverse direction. The flying wing  10  may have a lifting body  12  and a propulsion element  14 . The entire lifting body  12  may act as a wing to provide lift to the flying wing  10 . The flying wing  10  optionally may include one or more internal compartments such as a passenger compartment, a cockpit and a fuel tank. A windshield  16  may be located in the front transverse region  18 , where a cockpit may be housed. Similarly, the windows  20  may provide a view to persons in a passenger compartment. The flying wing  10  may also include additional compartments for storing cargo or other materials. 
     The propulsion element  14  may be rotatably attached to the lifting body  12  and may comprise a platform and one or more, for example two, jet engines  22 . The propulsion element  14  may optionally have vertical stabilizers. When the flying wing  10  is flown at subsonic speeds, the propulsion element  14  may be positioned at a yaw angle such that the engines  22  provide thrust in a direction parallel to the transverse axis of the lifting body  12  such that the front transverse region  18  may face the direction of travel and the back transverse region  24  may be in the center of the trailing edge of the lifting body  12 . The lifting body  12  may also be rotated 90°, or any angle, in the yaw direction and positioned to alter the direction of thrust such that the flying wing  10  flies in a direction parallel to a longitudinal axis such that the front longitudinal region  26  faces the direction of travel and the back longitudinal region  28  may be at the center of the trailing edge of the lifting body  12 . 
     In this embodiment, the two jet engines  22  may be mounted on the propulsion element  14 . However, other propulsion mechanisms, for example rocket engines, may optionally be used to provide thrust. The propulsion element  14  in this embodiment may have a circular platform. Optionally, the propulsion element  14  may have the shape of a ring, an ellipse, a polygon or other shape. Optionally, more than one propulsion element  14  may be mounted on the lifting body  12 , each supporting one or more engines. The propulsion element  14  may also optionally include one or more vertical stabilizers. 
     The propulsion element  14  may be mounted at any location, for example in the center, of the upper surface of the lifting body  12  to generate sufficient longitudinal stability margin by adjusting the location of center of gravity. The propulsion element  14  may optionally be mounted on the lower surface of the lifting body  12  or the lifting body  12  may have one or more propulsion elements  14  on both its upper and lower surfaces. The propulsion element  14  may also be optionally mounted on tracks that allow the propulsion element  14  to rotate and to move to a different location on the lifting body  12 . Additional propulsion elements  14  may also optionally be used. 
     Now referring to  FIG. 2 , the planform  30  of the flying wing  10  of  FIG. 1  having the lifting body  12  rotated 90°. As used herein, “planform” refers to the image of an aircraft viewed from above. The distance from front transverse tip  32  to back transverse tip  34  is the width of the lifting body  12  and is substantially less than the length of the lifting body  12 , measured from the front longitudinal tip  36  to the back longitudinal tip  38 . In this embodiment, the length may be approximately 40 meters, but may optionally be longer or shorter, depending in part on the mission requirements. 
     Lifting body  12  has double bilateral symmetry. That is, planform  30  is bilaterally symmetric across both longitudinal axis  40  and transverse axis  42 . The flying wing  10  tapers toward the front longitudinal tip  36  and the back longitudinal tip  38  such that the flying wing  10  has a high sweep angle during flight in a direction parallel to the longitudinal axis. Conversely, when the flying wing  10  travels in a direction parallel to the transverse axis, it has a low sweep angle. As used herein, “sweep angle” refers to the angle between the leading edge of the wing and a line perpendicular to the direction of flight. 
     The flying wing  10  has a high aspect ratio when flown in a direction parallel to the transverse axis  42 , and a low aspect ratio when flown in a direction parallel to the longitudinal axis  40 . As used herein, “aspect ratio” refers to the square of the wingspan divided by the planform area, AR=b 2 /S. Because the entire lifting body  12  acts as an airfoil wing, the aspect ratio during flight parallel to the longitudinal axis  40  is equal to square of the width, the distance between the front transverse tip  32  and the back transverse tip  34 , divided by the planform area. The aspect ratio during flight parallel to the transverse axis  42  is equal to the square of the width, the distance between the front longitudinal tip  36  and the back longitudinal tip  38 , divided by the same planform area as flying in the direction of the longitudinal axis. 
     When flying in a direction parallel to the longitudinal axis  40 , the front longitudinal region  26  may be positioned at the front of the flying wing  10  and the back longitudinal region  28  may be positioned at the back of the flying wing  10 . In this configuration, the leading edge of the planform  30  may include the right edge  44  and the left leading edge  46 , both having a high sweep angle when the flying wing is flown in parallel to the longitudinal axis. In this configuration, the trailing edge of the planform  30  includes the right edge  48  and the left edge  50 . The high sweep angle may reduce wave drag and increase efficiency at supersonic speeds. Generally the sweep angle of the edges  44  and  46  during longitudinal flight may be between 40° and 88°, and may be about 60°. 
     Optionally, leading and trailing edges of a planform may vary along the length of the leading and trailing edges, as they do in planform  30 . In planform  30 , the sweep angle may be greatest at the front longitudinal region  26 , where the longitudinal right edge  52  and the longitudinal left edge  54  may have sweep angles of generally about 60° or greater, and may be about 80°. 
     The back longitudinal region  28 , being symmetrical with the front longitudinal region  26 , may include the negatively swept back right edge  56  and the back left edge  58 . These very high sweep angle regions may create air effects similar to the effects resulting from leading air extensions or strakes as incorporated in various supersonic aircraft, for example the F/A-18 Hornet. This high sweep angle in the longitudinal direction may reduce wave drag and increase fuel efficiency during supersonic flight in a direction parallel to the longitudinal axis. 
     The right central edge  60  and the left central edge  62  may be part of the leading edge when the flying wing  10  flies in a direction parallel to the longitudinal axis. This may provide a sweep angle of about 60°, and may be less than the sweep angle of the front longitudinal region  26 . The Corresponding right central edge  64  and the corresponding left central edge  66  may have high negative sweep angles due to their symmetry with the right central edge  60  and the left central edge  62 , respectively. 
     The front transverse region  18  and the back transverse region  24  may include the longitudinal edge  68  and the longitudinal edge  70 , respectively, that may each have sweep angles of about 45°, and may be less than the sweep angle of the right central edge  60  and the left central edge  62 . Likewise, the front transverse edge  72  and the back transverse edge  74  are mirror images of corresponding front transverse edge  68  and back transverse edge  70 . 
     When the flying wing  10  travels in a direction parallel to the transverse axis  42 , the right edge  48  become the right leading edge, the right edge  44  becomes the left leading edge, the left edge  46  becomes the left trailing edge and the left edge  50  becomes the right trailing edge. In this configuration, the sweep angles of the edges  44  and  48  are equal to the complimentary angles of their sweep angles in the configuration for longitudinal flight. Thus, edge  52  of the front longitudinal region, having a longitudinal sweep angle of about 60° or more, has a transverse sweep angle of about 30° or less. The edge  56  of back longitudinal region  28 , being symmetric, also has a transverse sweep angle of 30° or less. Edges  68  and  72  of the front transverse region are about 45°, and the central edges  60  and  64  are between about 30°. Therefore, during flight in a direction parallel to the transverse axis, the flying wing  10  has a low sweep angle. 
     Referring now to  FIG. 3 , an alternative planform  80  has a blunted front transverse region  82  and a blunted back transverse region  84 , different from the pointed front transverse region  18  and back transverse region  24  shown in  FIGS. 1 and 2 . The Planform  80  also has a left edge  86  and a right edges  88  that have a left central edge  90  and a right central edge  92 , both having sweep angles between 45° and 60° during flight parallel to a transverse axis. The left front longitudinal edge  94  and the right front longitudinal edge  96  each may have a sweep angle of about 60°. The front longitudinal tip  98  may have a very high sweep angle of about 70° or greater, and may be about 80°. The front longitudinal tip  98  may be pivotably attached at a pivot point  100 . 
     The left edge  102  and the right edge  104  are mirror images of the left edge  86  and the right edge  88 . The back longitudinal tip  106  may have a sweep angle of about 70° or greater and may be pivotably attached at pivot point  108 . 
     Referring now to  FIG. 4 , during flight parallel to a transverse axis, the front longitudinal tip  109  and the back longitudinal tip  111  may pivot upward such that they are substantially vertical, approximately 90° relative to the flying wing  110 . In this embodiment, the front longitudinal tip  109  and the back longitudinal tip  111  may be pointed. Optionally, the front longitudinal tip  109  and the back longitudinal tip  111  may be rounded, blunted or squared off and may also optionally fold downward, or both. Allowing the front longitudinal tip  106  and the back longitudinal tip  108  to pivot to a vertical position may provide stability during subsonic flight. 
     Alternative planforms may be provided in accordance with the invention. The lengths of planforms may be substantially greater than the widths such that during flight parallel to a longitudinal axis the planform may have an aspect ratio between 0.1 and 5 and during flight parallel to a transverse axis the planform may have an aspect ratio between 1 and 20. 
     Referring now to  FIG. 5 , the longitudinal cross-section of the flying wing  10  shows the supersonic airfoil  152  at the centerline, i.e. at the longitudinal plane  40  in  FIG. 2 . The airfoil  152  may have an isentropic compression lower surface  154  which may be completely flat. As used herein, “isentropic” refers to a compression surface generating very weak or no shock waves. At a 0° angle of attack, a flat surface may be the most isentropic surface. This isentropic compression lower surface  154  may minimize the sonic boom, which may be as low as 0.3 psf. on the ground below the flying wing  10 . As used herein, “angle of attack” refers to the angle between an airfoil and the direction of flight. Optionally, the lower surface  154  may be slightly curved, and may include a slight negative camber, such that the flying wing  10  may have an isentropic compression lower surface when at an angle of attack slightly above 0°, for example at a 2° angle of attack. As used herein the term “camber” refers to the curvature of the airfoil. 
     The upper surface  156  may be curved, having a camber that gives it a height  158  and may contribute to providing lift. The three dimensional shape of the aircraft, having a flat, isentropic compression lower surface  154  and upper camber creating a curved upper surface  156  and pointed longitudinal tips  36  and  38  may give the flying wing  10  a positive lift to drag ratio at a 0° angle of attack. In this embodiment, the upper surface  156  may have a constant curvature and may be defined by a single radius of curvature. The upper surface  156  may optionally be substantially flat in the center or may have varying degrees of curvature across the length. 
     Referring now to  FIG. 6 , the transverse cross section of the aircraft  10  shows the subsonic airfoil  160 . The lower surface  154  may be flat when viewed from this direction also. The upper surface of the subsonic airfoil  160  may have a convex curvature similar to the camber of the supersonic airfoil  152 . The subsonic airfoil  160 , however, may include an outer region  162  and a central region  164 . The outer region  162  may have a slight curvature or may be substantially straight from the transverse edges  32  and  34  up to the central region  164 . The central region  164  may have a more pronounced curvature, raising it to accommodate a passenger, storage or other compartment or other internal features. The subsonic airfoil  160  may be thicker, having more camber than the relatively thin airfoil  152  used during supersonic flight. Thus, the flying wing  10  may have a relatively thin supersonic airfoil shown in  FIG. 5  and a relatively thick subsonic airfoil shown in  FIG. 6 . 
     All of the airfoil shown in  FIGS. 5 and 6  has an isentropic compression (flat at 0° angle of attack) lower surface. Flat lower surfaces may be isentropic at a 0° angle of attack and may produce a very weak or no downward shockwave. Optionally, the lower surface may be cambered, having a slightly curved surface, either concave or convex, to provide greater lift and/or to minimize the shock wave and associated sonic boom at a certain angle of attack. 
     The leading and trailing edges of airfoils in the  FIGS. 1-6  are depicted as being substantially sharp edges. Supersonic airfoils may generally include sharp leading and trailing edges to improve performance at high speed. Subsonic airfoils may perform better when utilizing a blunted leading edge. In addition, the supersonic airfoil may optionally utilize slightly blunted leading edges. A blunted leading or trailing edge may provide air affects over the surface of the flying wing that increase the critical angle. As used herein the “critical angle” is the angle of attack at which an aircraft stalls. A sharp leading edge typically decreases the critical angle of an airfoil, while a blunt leading edge typically increases the critical angle. 
     Referring now to  FIG. 7 , an edge of an airfoil  220  may house a slat  222 . The slat  222  may be extended from a leading edge to create a blunt edge. This may provide more lift and/or increase the critical angle of attack and/or increase the optimal angle. The slat  222  is shown extending from an airfoil edge  220  in a downward direction. The slat  222  may generally be included on any edges of a flying wing used as leading edges during flight in either direction. For example, the edges  44 ,  46  and  48  shown in  FIG. 2  all serve as leading edges and may include slats. The slats may optionally be included only in edges that serve as leading edges during subsonic flight in the transverse direction, for example edges  44  and  48  in  FIG. 2 , or may optionally also be included in leading edges during supersonic flight, for example edge  46  in  FIG. 2 . When the aerodynamic affects caused by the slat are desired, it may be extended as shown in  FIG. 7 . Otherwise, the slat  222  may be housed inside the airfoil edge  220 . 
     Referring now to  FIG. 8 , an alternative method of providing bluntness to an airfoil edge  224  uses a radial air ejector  226 . Radial air ejector  226  may project high pressure airflow radially in multiple directions  228  from a leading edge  224 . As shown in  FIG. 9 , a radial airflow  230  may form a blunt edge  232  when it encounters a unidirectional airflow  234 . The amount of bluntness may be adjusted by adjusting the amount of radial airflow ejected from the radial air flow ejector  226 . This method of creating a virtual blunt edge on an airfoil may require fewer moving parts and may add less weight to the overall design, and may thereby increase efficiency. 
     It will be appreciated by persons skilled in the art that the present invention is not limited to what has been particularly shown and described herein above. In addition, unless mention was made above to the contrary, it should be noted that all of the accompanying drawings are not to scale. A variety of modifications and variations are possible in light of the above teachings without departing from the scope and spirit of the invention, which is limited only by the following claims.