Patent Publication Number: US-2007114327-A1

Title: Wing load alleviation apparatus and method

Description:
FIELD OF THE INVENTION  
      The present invention relates to aircraft, and more particularly to a system adapted to alleviate lift-induced structural-bending loads experienced by the wings of an aircraft during flight.  
     BACKGROUND OF THE INVENTION  
      The wing structure of a typical, modern day jet aircraft is designed at least in part by considering critical loads at limiting flight or ground conditions. Typically, a limiting flight condition is one at which high load factors are experienced, and is one that is usually avoided during normal flight operations. The wing structure has to be designed with sufficient strength to thus be able to accommodate the high load factors that are experienced at a limiting flight condition, even though such a condition will rarely, or possibly never, be encountered during flight of the aircraft.  
      Designing wing structure to accommodate the high load factors that are experienced at limiting flight conditions requires that the wing spars and other structural components within the wing be made sufficiently robust to withstand the high load factors. However, this results in a wing that is heavier than would otherwise be required to accommodate normal load factors that are typically experienced during flight.  
      The overall structural weight of the wing and/or attachment structure for attaching the wings to the fuselage could be reduced if at key critical load conditions the spanwise location of the lift experienced by each wing was to be moved more inboard and/or reduced in magnitude during flight. Reducing the overall weight of the wings would result in a lighter aircraft that is able to fly further with a given payload. Alternatively, moving the spanwise aerodynamic load distribution more inboard along the wings, would allow the aircraft to accommodate even more revenue-generating payload, thus enhancing the value of the aircraft. Being able to move the lift-inducing structural-bending forces experienced by the wings more inboard towards the fuselage of the aircraft would also allow the wing span of the wings to be increased while retaining much of the original wing frame and attachment structure (i.e., with less structural weight for the extended length wings).  
     SUMMARY OF THE INVENTION  
      The present invention is directed to an apparatus and method for alleviating the lift-inducing forces experienced near the outer tips of the wings of an aircraft, and moving the lift-inducing structural bending forces more inboard, spanwise, along the wings towards the fuselage of the aircraft.  
      In one preferred embodiment, deployable panels are located in an upper surface of each wing at a spanwise location that is at least about halfway out to the tip of the wing. When the panels are deployed into the airstream, this reduces the local aerodynamic loads experienced at the outer tips of the wings and effectively moves the bending forces more inboard (i.e., spanwise) along the wings towards the fuselage. The panels, in one preferred form, are deployed by actuators mounted within each wing. The actuators are in turn controlled by a flight control system on the aircraft.  
      By reducing the aerodynamic load distribution experienced at the outboard half of the wings, and effectively moving this force more inboard along the wings closer to the fuselage, the maximum payload able to be carried by the aircraft can be increased. The aerodynamic load induced bending moment on the wing is defined as follows:  
         M   ⁡     (     γ   0     )       =       ∫     γ   0       b   /   2       ⁢       1   2     ⁢   ρ   ⁢           ⁢     v   2     ⁢       C   L     ⁡     (   γ   )       ⁢     (     γ   -     γ   0       )     ⁢     c   ⁡     (   γ   )       ⁢     ⅆ   γ             
          where: γ is a spanwise distance coordinate; γ 0  is a particular spanwise location; M(γ 0 ) is the aerodynamic load induced bending moment on the wing at spanwise coordinate γ 0 ; C L (γ) is lift coefficient at spanwise coordinate γ; ρ is air density; ν is airspeed; c(γ) is wing chord at spanwise coordinate γ; and b/2 is the semispan of the aircraft. 
 
 Alternatively, the internal structure of the wings (e.g., wing spars) can be made lighter in weight because of the reduced aerodynamic loads and induced bending moments that need to be accommodated by the wings. Alternatively, longer wings could be employed without requiring significantly heavier structure. 
       

      Further areas of applicability of the present invention will become apparent from the detailed description provided hereinafter. It should be understood that the detailed description and specific examples, while indicating various preferred embodiments of the invention, are intended for purposes of illustration only and are not intended to limit the scope of the invention. 
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS  
      The present invention will become more fully understood from the detailed description and the accompanying drawings, wherein:  
       FIGS. 1A, 1B  and  1 C are plan views of aircraft incorporating preferred embodiments of the wing load alleviation device in each wing;  
       FIG. 2  is an enlarged perspective view of an outermost portion of one of the wings of the aircraft in  FIG. 1A  showing a panel of the wing load alleviation device in solid lines in its retracted position;  
       FIG. 3  is a simplified side cross sectional view of the wing of the aircraft in  FIG. 1A  taken in accordance with section line  3 - 3  in  FIG. 1A  illustrating the panel in the retracted position;  
       FIG. 4  is a view of the wing of  FIG. 3 , but showing the panel in the deployed (i.e. extended) position; and  
       FIG. 5  is a graph illustrating the load experienced by one of the wings over its spanwise length with the panel deployed, and also with the panel in its retracted position. 
    
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS  
      The following description of the preferred embodiment(s) is merely exemplary in nature and is in no way intended to limit the invention, its application, or uses.  
      Referring to  FIG. 1A , a wing load alleviation apparatus  10  in accordance with a preferred embodiment of the present invention is illustrated located in each wing  12  of an aircraft  14 . In this example the aircraft is a modern day, commercial jet aircraft having a flight control system  15 , although it will be appreciated that the apparatus  10  could be employed in propeller or turboprop driven aircraft as well. The aircraft may be a subsonic transport equipped with a swept, moderate or high aspect ratio wing and turbofan engines. The wing could employ metallic structure such as structure using aluminum alloy material, or composite structure such as structure using carbon-epoxy or other composite material, or a hybrid of metallic and composite structure. The apparatus  10  may be located at any outboard spanwise point along the length of its associated wing  12 , although preferably it is located at a spanwise position near to the tip  16  of its associated wing  12  and, more preferably, at least about halfway out of the tip  16 . Even more preferably, the apparatus  10  is positioned outboardly, of an outboard-most trailing edge device, as will be explained more fully in the following paragraphs.  
      The apparatus  10  is in communication with a command generator  15 B for generating commands to control the apparatus. A sensor system  15 A is used to sense the presence of lift-inducing structural-bending forces and moments being experienced by the wings. The sensor system  15 A may comprise one or more of an inertial load factor sensor, a pitch rate sensor or even strain gauge sensors  15 A, positioned in each wing  12 . The command generator  15 B may comprise a microprocessor, a digital computer, an analog computer, or any other form of system capable of generating the required command signals to the apparatus  10 . The command generator  15 B is able to apply commands to each apparatus  10  in each wing  12  independently if needed. Commands may be based on information from the sensor system  15 A, the flight control system  15 , or a combination of both, as well as from pilot input(s). The apparatus  10  could also be controlled in conjunction with other deployable components on the wing  12 , such as ailerons, flaperons, elevons, flaps, etc.  
      In the embodiment of  FIG. 1A , the air deflecting members forming the apparatus  10  are located outboard of the outboard-most of the ailerons  42 , as well as outboard of the trailing edge flaps  40  and the typical spoilers or speed brake panels  41 .  
       FIG. 1B  illustrates a variant preferred embodiment in which the outboard-most trailing edge devices comprise the outboard-most pair of flaps  40 , and where a pair of air deflecting panels forming the apparatus  10  are located at least in part outboard and forward of the outboard-most pair of flaps  40 . The embodiment of  FIG. 1B  also shows the apparatus  10  as including another pair of air deflecting panels located on the inner surfaces of upwardly-oriented winglet members  43  at the outer ends of the wings  12 .  
       FIG. 1C  illustrates another preferred embodiment in which upper edges of air deflecting panels  10 U are adjacent to the upper surface of the wings  12 , at least in part outboard of the outboard-most pair of flaps  40  and, in this case, forward of a rear spar  32 . In this embodiment, the air deflecting panels of the apparatus  10  can be controllably extended by at least one activator (inside the wing  12  and so not shown) translating the upper edge upward into the airstream above the wing  12 .  FIG. 1C  shows panels  10 U forward of the rear spar  32 ; however, an alternate embodiment would place them aft of the rear spar  32 .  
      With reference to  FIGS. 2-4 , the apparatus  10  includes a panel  20  that is movable by an actuator  22  mounted within the wing  12 . The panel  20  could be a flap or spoiler. The panel  20  includes a leading edge  20   b  and a trailing edge  20   c . The panel is coupled to the wing  12  for pivotal movement about or near its leading edge  20   b . The actuator may be an electrical, hydraulic, electrohydraulic or pneumatic actuator, or any other suitable form of actuator that is able to move the panel  20  into an airstream  24  flowing over the wing  12 . Many forms of actuators that are suitable for use in aircraft and aerospace applications could be employed. The actuator  22  is coupled with the flight control system  15 .  
      With specific reference to  FIG. 2 , the wing  12  is otherwise of conventional construction and typically includes one or more leading edge slats  26  and one or more ailerons  28 . A front spar  30  and a rear spar  32  are indicated in dash lines, as is a wing tip spar box  34 . Thus, the inclusion of the wing load alleviation device  10  does not otherwise require significant alternation of the traditional construction and internal components of the wing  12 . From  FIGS. 2 and 3 , it can be seen that the panel  20  includes an outer surface  20   a  that has a contour generally in accordance with a contour of an upper surface  12   a  of the wing  12 . Thus, when the panel  20  is in its retracted position as shown in  FIG. 3 , the panel has no tangible affect on the airflow  24  over the wing  12 . The panel  20  may be made of aluminum, from composites, or from other suitably strong, lightweight and durable materials. The panel  20  is also preferably located outboardly, spanwise, of the aileron  28 , which in this example is the outboard-most trailing edge device on the wing  12 . It will be appreciated that the outboard-most trailing edge device could also be a flap, a flaperon or an elevon. The panel  20  is also preferably located rearwardly of the rear spar  32  and elevationally above the midplane of the wing box formed by the spars  30  and  32  and upper and lower wing skins.  
      With continued reference to  FIGS. 3 and 4 , when the actuator  22  receives a signal from the command generator  15 B to deploy the panel  20 , the actuator extends the panel  20  from the stowed position shown in  FIG. 3  into the deployed position shown in  FIG. 4 . This has the effect of reducing local aerodynamic lift forces near the panel  20 , and thus reducing the aerodynamic lift induced load distribution experienced along the wing inboard to the side of body. The aerodynamic lift-induced load distribution is effectively shifted spanwise along the wing  12  towards the centerline (C L  in  FIG. 1 ) of the fuselage  14   a  of the aircraft  14 . Moving the lift-inducing structural-bending forces and moments more inboard spanwise along the wings  12  has the effect on wing loads of equivalently decreasing the wing span.  
      By effectively shifting the aerodynamic load distribution more inboard, spanwise, along the wings  12 , the apparatus  10  allows an even greater payload to be carried by the aircraft  12  than what would otherwise be possible without the use of the apparatus  10 . Alternatively, the structural framework of the wings  12  could be made lighter in weight because the maximum aerodynamic loads that each wing needs to be able to accommodate would be less when the apparatus  10  is employed in each wing  12 . Still further, the use of the apparatus  10  in each wing  12  could alternatively allow a wing of even longer span to be used with a given wing structural design.  
      Referring to  FIG. 5 , a conceptual graph of the spanwise load experienced by the wing at various spanwise locations along the wing is illustrated when the apparatus  10  is deployed and also when it is retracted. Curve  36  represents the spanwise load with the apparatus  10  in its deployed position, while curve  38  represents the spanwise load with the apparatus  10  in its retracted position. Curve  36  illustrates that a greater portion of the load is shifted toward the center line (C L ) of the fuselage  14   a  of the aircraft  14  with the apparatus  10  in its deployed position. Curve  36  also indicates that the spanwise load experienced near the wing tip  16  is significantly reduced with the apparatus  10  deployed. A wing load alleviation control law (i.e. algorithm) will typically command deployment of the panel  20  when the aircraft is experiencing a current or anticipated maneuver or gust load factor above a threshold value, which may be at least or near a limit load factor (2.5 g&#39;s), or more generally anywhere between 1.25 g&#39;s and 3.75 g&#39;s. The current load factor can be obtained from an inertial load sensor sensing load factor N z , and the anticipated load factor can be synthesized from a combination of N z  and at least one of {dot over (N)} z  and pitch acceleration {dot over (q)}. Individual left and right wing load factor connections may be completed as a function of roll acceleration {dot over (p)}, in one preferred embodiment.  
      The apparatus  10  can be employed on virtually any form of airborne mobile platform that makes use of wings. The apparatus can be used in connection with wings having a winglet, a wing tip extension, or both, or a raked tip. In such instances, the panel  20  could be located inboardly of the winglet, wing tip extension or raked tip, or possibly within a portion of the wing tip extension or raked tip. In such instances, the downward incremental life generated by the panel  20  may be enhanced by the presence of the winglet, wing tip extension or raked tip. The command generator  15 B can be used in connection with a suitable algorithm to apply suitable control signals to each apparatus  10  independently of the other and also in response to the detection of a maneuver limit load condition being exceeded, or about to be exceeded, or the detection of the actual or incipient detection of the exceedance of gust limit load conditions.  
      Furthermore, the apparatus  10  can be employed on wings that are formed with aluminum, composite materials, etc., and therefore is not limited to any specific material construction that is employed on the wings.  
      While various preferred embodiments have been described, those skilled in the art will recognize modifications or variations which might be made without departing from the inventive concept. The examples illustrate the invention and are not intended to limit it. Therefore, the description and claims should be interpreted liberally with only such limitation as is necessary in view of the pertinent prior art.