Abstract:
A prismatic vortex generator for attenuating flow separation which occurs during supersonic flow of air over structure such as an aircraft airfoil, its fuselage, surfaces forming a part of a jet engine inlet, or similar surfaces subjected to supersonic airflow. A series of prismatic vortex generators are provided, each of which is configured to generate a vortex which attenuates flow separation and weight drag resulting from the supersonic airflow. Each prismatic vortex generator has a prismatic shape with a base, leading and trailing ends, and sidewalls that incline toward and join each other to form an apex. The leading end of each prismatic vortex generators is inclined away from the direction of flow.

Description:
FIELD OF THE INVENTION 
       [0001]    The present invention relates in general to vortex generators for supersonic aircraft and, in particular, to increasing the robustness of vortex generators for maintaining an attached boundary layer in supersonic aircraft inlet engine applications. 
       BACKGROUND OF THE INVENTION 
       [0002]    Supersonic airflow over components such as the internal portion of aircraft engine inlets and aircraft airfoils, can generate shock waves. These shock waves oscillate back and forth and causes the air flow to separate from the surface within engine inlets or on airfoils. Flow separation on airfoils can result in loss of lift and can ultimately cause a stall. Further, these shock waves can cause increased drag and buffeting of control surfaces attached to the trailing edge of the wing. Similar separation conditions can also occur at engine inlets, and other areas of the aircraft. 
         [0003]    One approach to preventing or attenuating flow separation is through the introduction of vortices in the boundary layer or sub-boundary layer region. Vortex generators that project from the surface of the engine inlet or airfoil in supersonic airflow applications can thus be utilized to prevent or attenuate flow separation. The vortex generators proposed previously for supersonic engine inlets primarily consist of a plurality of short (smaller than the boundary layer height), thin rectangular blades (or microvanes) located downstream of the leading edge of the engine inlet and upstream of the point on the surface within the engine inlet where flow separation would occur without the vortex generator. 
         [0004]    In a typical supersonic application, the blades are generally aligned with the path of air or at an acute angle with respect to the line of flight. Vortices that form off the tips of the blades can draw air down from the upper region of the boundary layer down toward the near-wall region to attenuate flow separation and reduce unwanted aerodynamic phenomena associated with flow separation, such as drag, aerodynamic blockage, and shock dithering. However, rectangular vortex generators are deficient for practical applications when considering needs for reliability, maintainability, and survivability particularly in engine inlets. Thin, rectangular vortex generators lack the robustness to survive the harsh pressure and aerodynamic loads, and heat erosion. The thin, rectangular vortex generators also do not tolerate damage due to debris that may impact the vortex generator. Furthermore, rectangular microvanes have been historically difficult to incorporate into a producible manufacturing process. 
         [0005]    While current state-of-the-art vortex generator designs and integration approaches are feasible for attenuating flow separation in supersonic applications, an improved solution would be desirable to increase the robustness of vortex generators. 
       SUMMARY OF THE INVENTION 
       [0006]    One embodiment for a vortex generating system for preventing flow separation in supersonic airflow over a surface comprises a plurality of prismatic-shaped vortex generators or structures attached to the airflow surface. These new vortex generators are also referred to as prismatic-shaped passive flow control devices. The general shape of these passive flow control devices in one embodiment is prismatic, having a triangular longitudinal cross-section, trapezoidal side view, and rectangular base. Other embodiments of passive flow control devices can have a hexangular base. The present invention addresses the problems described above in the background. 
         [0007]    The prismatic vortex generators can be located on a surface inside an aircraft engine inlet and located in a region downstream of the engine inlet intake. The prismatic vortex generators are further located upstream on the surface where flow separation would occur without the prismatic vortex generator. The vortices generated by the prismatic vortex generators draw air down from the upper region of the boundary layer down toward the near-wall region to attenuate flow separation that can negatively affect engine performance. 
         [0008]    Depending on the application, the prismatic vortex generators can be oriented such that they are at an angle with respect to the direction of flow. The spacing between the prismatic vortex generators can also vary by application. 
         [0009]    The prismatic shape of the vortex generators makes them more robust to thus advantageously increase their reliability, maintainability, and survivability. In this embodiment, the inclined or slanted triangular faces at each end of a prismatic vortex generator allow the prismatic vortex generator to be more resistant to harsh pressure and aerodynamic loads, and erosion due to heat. In this embodiment, a pair of side walls join the inclined triangular faces. The sidewalls, like the triangular faces, are inclined such that they meet to form an apex that runs from the apex of one triangular face to the other. Each of the sidewalls in this example thus has a trapezoidal shape. The faces and sidewalls are joined at the bottom by a base. The inclined triangular faces and inclined sidewalls allow the prismatic vortex generator to better tolerate impact from debris by allowing debris to deflect of the inclined surfaces rather than experience a direct impact. 
         [0010]    The features of prismatic-shaped vortex generators thus provide several improvements for vortex generators. Use of the prismatic vortex generators is not restricted to the engine inlet supersonic aircraft as they may be used on other supersonic airflow surfaces such as a wing structure, fuselage, or the tail airfoils of an aircraft, or other structures subjected to supersonic airflow such as turbine or compressor blades or the like. Other applications of prismatic vortex generators include improving the health of the boundary layer, mitigating the growth of undesired vortices, stabilizing the position of oblique and normal shockwaves, and enhancing mixing between multiple streams or regions of flow. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0011]      FIG. 1  is a sectional side view of a conventional, aircraft engine inlet equipped with a conventional rectangular vortex generator; 
           [0012]      FIGS. 2A-2C  show isometric, forward looking aft, and top views of the prior art rectangular vortex generators; 
           [0013]      FIGS. 3A-3C  show isometric, forward looking aft, and top views of one embodiment of an aircraft engine inlet with prismatic-shaped vortex generators, constructed in accordance with the present invention; 
           [0014]      FIG. 4  is a sectional side view showing the location of the prismatic-shaped vortex generator within an aircraft engine inlet, constructed in accordance with the present invention; 
           [0015]      FIGS. 5A-5C  show schematic side, front and top views of an individual prismatic-shaped vortex generator of an embodiment constructed in accordance with the present invention; 
           [0016]      FIGS. 6A-6C  show schematic side, front and top views of an individual prismatic-shaped vortex generator of an embodiment constructed in accordance with the present invention. 
       
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       [0017]      FIG. 1  shows a conventional engine inlet or duct  10  with an inner surface  14  over which air can flow. A boundary layer  16  with a measurable thickness develops over the surface  14  as air flows over the surface  14 . When air flows over the surface  14  at supersonic speeds, shock waves  18  can be generated. The shock waves  18  oscillate back and forth inside the engine inlet  10  and can cause the boundary layer  16  to separate from the surface  14  of the engine inlets  10 . To attenuate this flow separation of the boundary layer, a conventional rectangular vortex generator  20  can be located on the surface  14 . The rectangular vortex generator  20  creates vortices than can help prevent flow separation. 
         [0018]    As illustrated in  FIGS. 2A-2C , a series of conventional, rectangular vortex generators  20  can be located on the surface  14  and can be oriented at an angle relative to the direction of air flow. Although, this conventional arrangement of rectangular vortex generators may attenuate flow separation at supersonic conditions, the rectangular vortex generators  20  are not robust enough for the reliability, maintainability, and survivability desired as the rectangular faces  22  of the vortex generators  20  do not effectively tolerate the harsh pressure and aerodynamic loads, heat erosion, debris impact nor are these devices producible with standard manufacturing processes. 
         [0019]    Referring to  FIGS. 3A-3C  an embodiment of a vortex generating system is shown comprising prismatic-shaped vortex generators or structures  40  attached to the surface  14  of the engine inlet  10 . In this embodiment, the prismatic-shaped vortex generators  40  are shown spaced apart and with an orientation such that the prismatic vortex generators  40  are at an angle relative to the direction of air flow, i.e. angle of attack “AoA”. Both the spacing and angle of attack is dependant on the application. In this embodiment, the spacing can be determined by the ratio between a distance “D” between a pair of prismatic vortex generators  40  and a distance “d” between the individual generators that form a pair. The ratio of D/d is preferably 2 but can vary in a range between 1.5 and 2.5. The AoA of a prismatic vortex generators  40  is preferably 20 degrees but can vary between 16 and 24 degrees. 
         [0020]    In this embodiment, the prismatic-shaped vortex generators  40  can be located within an engine inlet  10  and located in a region downstream of the engine inlet  10  intake, but upstream from the throat or minimum cross-sectional flow area, as shown in  FIG. 4 .  FIG. 4  further shows the prismatic-shaped vortex generators  40  located upstream on the surface where flow separation would occur without the vortex generator. This streamwise location is preferably located in the range of 8 to 10 times the thickness of the boundary layer  16  upstream of the shockwave intersection with the boundary layer edge. Vortices generated by the prismatic vortex generators  40  draw air down from the upper region of the boundary layer  16  down toward the near-wall region of the airflow surface  14  to help prevent and attenuate flow separation that can occur due to shock waves  18  generated during supersonic air flow, thus helping prevent negative effects on engine performance. 
         [0021]    Referring to  FIGS. 5A-5C , the geometry of a prismatic-shaped vortex generator  40  is shown in more detail. In this embodiment, the prismatic-shaped vortex generator  40  comprises a base  42  having a rectangular shape, and inclined polygonally-shaped sidewalls  44  joined at their ends by inclined polygonally-shaped ends  46  and  47 . The prismatic vortex generator  40  may be fabricated from a single piece of suitable material or multiple pieces joined or fastened together. The base  42  allows for attachment or fastening of the prismatic vortex generator  40  to the surface  14  of the engine inlet  10  ( FIG. 4 ). The base  42  preferably has a width of 1.5 times the height “H” of the prismatic vortex generator  40  but can vary between 1 to 2 times the height H. The inclined or slanted ends  46 ,  47  in this example are flat and incline toward each other. The inclined ends  46 ,  47  are preferably inclined 45 degrees from a vertical axis but the incline can range from 30 to 60 degrees. The upstream or leading end  46  slants in a downstream direction from base  42 . The downstream or trailing end  47  slants in an upstream direction from base  42 . Preferably ends  46 ,  47  are triangular in shape as shown in  FIG. 5B . However, other shapes for the ends  46 ,  47  such as trapezoidal can be utilized. The sidewalls  44  are flat and have a trapezoidal shape in this embodiment and like the triangular ends  46 ,  47 , are inclined relative to base  42 . Sidewalls  44  meet each other to form an apex  48  that runs from the apex of one triangular end  46  to the other end  47 , as best shown in  FIG. 5C . The inclined triangular ends  46 ,  47  and inclined sidewalls  44  allow the prismatic vortex generator  40  to better tolerate impact from debris by allowing debris to deflect of the inclined surfaces rather than experience a direct impact as well as better tolerate erosion due to viscous heating. 
         [0022]    The height H of the prismatic-shaped vortex generator  40  from base  42  to apex  48  is preferably in the range of ⅕ to ¼ the thickness of the boundary layer  16  but can vary between ⅕ to ½ the thickness of the boundary layer  16  ( FIG. 4 ). The height H remains constant from leading to trailing end  46 . Heights larger than ½ the thickness of the boundary layer may be undesirable as they can create undesirable increases in drag. The length “L” of the apex is preferably 8 times the height H but can vary between four and 12 times the height H depending on the application and the strength of the vortices required. The width of base  42  is shown to approximately be the same as the dimension of each sidewall  44  from base  42  to apex  48 . The shape in an end view as in  FIG. 5B  is thus that of a triangle. However, the width of base  42  may differ from the dimensions of sidewalls  44  from base  42  to apex  48 . 
         [0023]    In other embodiments, the prismatic-shaped vortex generators  40  could be disposed circumferentially spaced apart inside the engine inlet  10 . Thus, flow separation could be attenuated in all interior surfaces. 
         [0024]    In another embodiment, the prismatic-shaped vortex generators  40  could be located on an external surface such as the wing of an aircraft to prevent flow separation on the wing at supersonic speeds. 
         [0025]    In yet another embodiment, the faces  46  of the prismatic-shaped vortex generators  40  have a trapezoidal shape. In this embodiment, a flat top plate (not shown) would join the ends  46 ,  47  and sidewalls  44  to create a flat crest rather than a sharp apex  48 . 
         [0026]    In another embodiment shown in  FIGS. 6A-6C , a prismatic vortex generator  60  comprises a base  62  having a hexangular shape, and inclined polygonally-shaped sidewalls  64  joined at their ends by inclined polygonally-shaped ends  66  and  67 . Ends  66  and  67  are each comprised of two triangle-shaped sections having a side that coincides and form an end of an apex  68 . The two triangle-shaped sections making up the upstream or leading end  66  face away from each other and slant in a downstream direction from base  62  and also slant toward. The two triangle-shaped sections making up the downstream or trailing end  67  also face away from each other and slant in an upstream direction from base  62 . Sidewalls  64  meet each other to form an apex  68  that runs from the apex of one end  66  to the other end  67 , as best shown in  FIG. 6C . 
         [0027]    The invention has many advantages, including preventing boundary layer separation for supersonic applications and increasing the robustness of vortex generators to thus increase reliability, maintainability, and survivability of the vortex generator. It also provides increased resistance to heat erosion and debris impact. All of these advantages are provided simultaneously with a single device that is relatively easy to produce and adapt for use in air inlets. 
         [0028]    While the invention has been shown or described in only some of its forms, it should be apparent to those skilled in the art that it is not so limited, but is susceptible to various changes without departing from the scope of the invention.