Patent Document

BACKGROUND OF THE INVENTION 
       [0001]    The technology described herein relates generally to turbomachinery, particularly to gas turbine engines, and more particularly, to a gas turbine engine guide vane assembly. 
         [0002]    At least one known gas turbine engine assembly includes a fan assembly that is mounted upstream from a core gas turbine engine. During operation, a portion of the airflow discharged from the fan assembly is channeled downstream to the core gas turbine engine wherein the airflow is further compressed. The compressed airflow is then channeled into a combustor, mixed with fuel, and ignited to generate hot combustion gases. The combustion gases are then channeled to a turbine, which extracts energy from the combustion gases for powering the compressor, as well as producing useful work to propel an aircraft in flight. The other portion of the airflow discharged from the fan assembly exits the engine through a fan stream nozzle. 
         [0003]    To facilitate channeling the airflow from the fan assembly to the fan stream exhaust, at least one known gas turbine engine assembly includes an outlet guide vane assembly that is used to remove swirl before the fan nozzle. Such an outlet guide vane assembly is configured to turn the airflow discharged from the fan assembly to a substantially axial direction prior to the fan flow being exhausted from the bypass duct. In addition to turning the fan airflow, the outlet guide vane assembly also provides structural stiffness to the fan frame. More specifically, outlet guide vane assemblies generally include a plurality of outlet guide vanes that are coupled to the fan frame. 
         [0004]    In addition to outlet guide vanes, many fan frame assemblies include one or more (frequently two, diametrically opposed) dividing structures, often called “bifurcations”, which divide the annular space defined by the bypass duct into two semi-annular spaces. These dividing structures are typically hollow duct-like structures through which various mechanical, electrical, pneumatic, hydraulic, or other connections (including structural supports) can pass without causing disruption to the airflow through the bypass duct. The bifurcations “fair” or guide the flow in aerodynamic fashion around these structures, and may be integrated or blended into the profile of an upstream guide vane to reduce the number of individual airflow disruptions. 
         [0005]    Geometric sweep and lean characteristics for guide vanes have been previously demonstrated to be useful design parameters for reducing noise caused by aerodynamic interactions between guide vanes and upstream and/or downstream rotating elements such as fan blades. However, since bifurcations are typically radially oriented there remains a need for an improved approach to integrating advanced design swept and/or leaned guide vanes with bypass duct bifurcations. 
       BRIEF SUMMARY OF THE INVENTION 
       [0006]    In one aspect, an integrated outlet guide vane assembly for turbomachinery typically includes at least one outlet guide vane and at least one bifurcation having a leading edge and a trailing edge. The turbomachinery has a central axis of rotation and a defined direction of rotation about the axis. The guide vane comprises an airfoil having a leading edge and a trailing edge and has a non-zero angle of lean in the direction of rotation and a non-zero sweep angle relative to a line perpendicular to the central axis. The leading edge of the bifurcation has a non-zero angle of lean in the direction of rotation and a non-zero sweep angle relative to a line perpendicular to the central axis. The trailing edge of the vane is faired into the leading edge of the bifurcation. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0007]      FIG. 1  is a cross-sectional illustration of an exemplary gas turbine engine assembly; 
           [0008]      FIG. 2  is an elevational partial cross-sectional view of the gas turbine engine assembly shown in  FIG. 1 ; 
           [0009]      FIG. 3  is an elevational view of the guide vane assembly shown in  FIG. 2  taken along line  3 - 3 ; 
           [0010]      FIG. 4  is an illustration of the relationship in plan view between the fan assembly, guide vane assembly, and bifurcation shown in  FIG. 1 ; 
           [0011]      FIGS. 5 ,  6 , and  7  illustrate the relationship between guide vanes and the leading edge of the bifurcation shown in  FIG. 1 ; 
           [0012]      FIG. 8  is an elevational partial cross-sectional view similar to  FIG. 2  of another embodiment of the gas turbine engine assembly shown in  FIG. 1 ; and 
           [0013]      FIG. 9  is a perspective view of the gas turbine engine of  FIG. 8  in a typical installation configuration for an aircraft (not shown). 
       
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       [0014]      FIG. 1  is a cross-sectional schematic illustration of an exemplary gas turbine engine assembly  10  having a longitudinal axis  11 . Gas turbine engine assembly  10  includes a fan assembly  12  and a core gas turbine engine  13 . Core gas turbine engine  13  includes a high pressure compressor  14 , a combustor  16 , and a high pressure turbine  18 . In the exemplary embodiment, gas turbine engine assembly  10  also includes a low pressure turbine  20 , and a multi-stage booster compressor  22 , and a splitter  44  that substantially circumscribes booster  22 . 
         [0015]    Fan assembly  12  includes an array of fan blades  24  extending radially outward from a rotor disk  26 . Gas turbine engine assembly  10  has an intake side  28  and an exhaust side  30 . Fan assembly  12 , booster  22 , and turbine  20  are coupled together by a first rotor shaft  31 , and compressor  14  and turbine  18  are coupled together by a second rotor shaft  32 . 
         [0016]    In operation, air flows through fan assembly  12  and a first portion  50  of the airflow is channeled through booster  22 . The compressed air that is discharged from booster  22  is channeled through compressor  14  wherein the airflow is further compressed and delivered to combustor  16 . Hot products of combustion (not shown in  FIG. 1 ) from combustor  16  are utilized to drive turbines  18  and  20 , and turbine  20  is utilized to drive fan assembly  12  and booster  22  by way of shaft  31 . Gas turbine engine assembly  10  is operable at a range of operating conditions between design operating conditions and off-design operating conditions. 
         [0017]    A second portion  52  of the airflow discharged from fan assembly  12  is channeled through a bypass duct  40  to bypass a portion of the airflow from fan assembly  12  around the core gas turbine engine  13 . More specifically, bypass duct  40  extends between a fan casing or shroud  42  and splitter  44 . Accordingly, a first portion  50  of the airflow from fan assembly  12  is channeled through booster  22  and then into compressor  14  as described above, and a second portion  52  of the airflow from fan assembly  12  is channeled through bypass duct  40  to provide thrust for an aircraft, for example. Gas turbine engine assembly  10  also includes a fan frame assembly  60  to provide structural support for fan assembly  12  and is also utilized to couple fan assembly  12  to core gas turbine engine  13 . 
         [0018]    Fan frame assembly  60  includes a plurality of outlet guide vanes  70  that typically extend substantially radially, between a radially-outer mounting flange and a radially-inner mounting flange, and are circumferentially-spaced within bypass duct  40 . Fan frame assembly  60  may also include a plurality of struts that are coupled between a radially outer mounting flange and a radially inner mounting flange. In one embodiment, fan frame assembly  60  is fabricated in arcuate segments in which flanges are coupled to outlet guide vanes  70  and struts. In one embodiment, outlet guide vanes and struts are coupled coaxially within bypass duct  40 . Optionally, outlet guide vanes  70  may be coupled upstream or downstream from struts within bypass duct  40 . Guide vanes  70  serve to turn the airflow downstream from rotating blades such as fan blades  24 . 
         [0019]    Fan frame assembly  60  is one of various frame and support assemblies of gas turbine engine assembly  10  that are used to facilitate maintaining an orientation of various components within gas turbine engine assembly  10 . More specifically, such frame and support assemblies interconnect stationary components and provide rotor bearing supports. Fan frame assembly  60  is coupled downstream from fan assembly  12  within bypass duct  40  such that outlet guide vanes  70  and struts are circumferentially-spaced around the outlet of fan assembly  12  and extend across the airflow path discharged from fan assembly  12 . 
         [0020]      FIG. 1  also illustrates the bifurcations  80  and  82  which extend radially through the bypass duct  40  between the fan casing or shroud  42  and splitter  44 . The configuration of bifurcations  80  and  82  will be described in greater detail hereafter. While the figures herein illustrate two (upper and lower) bifurcations, it is possible that for certain configurations (including certain engine mounting arrangements) that either a single bifurcation or three or more bifurcations may be utilized. 
         [0021]      FIG. 2  is an enlarged elevational cross-sectional view of the gas turbine engine  10  of  FIG. 1 , showing the elements of  FIG. 1  in greater detail as well as illustrating the location at which the sectional elevational view of  FIG. 3  is taken along lines  3 - 3 . 
         [0022]      FIG. 3  is an elevational view which illustrates, looking rearward from the front of the gas turbine engine, the relationship of the vanes  70  to the reference lines and axes of the gas turbine engine  10 . As shown in  FIG. 3 , the guide vanes  70  are circumferentially distributed around the central axis  11  of the gas turbine engine  10 .  FIG. 3  illustrates the direction of rotation D of the gas turbine engine during normal operation, the radial direction R, and the lean angle L which leaned typical guide vanes  70  make with respect to the radial direction R. In the embodiment shown, the lean angle L shown in the direction of fan rotation, which provides maximum acoustic benefit. 
         [0023]      FIG. 4  is a plan view looking downward on elements of the gas turbine engine  10  to illustrate the relationship between the fan assembly  12 , guide vanes  70 , and bifurcation  80 . In this illustration, the guide vanes  70  incorporate lean in the direction of rotation, and a sweep angle toward the rear of the engine from their inner end (root)  72  to their outer end (tip)  74 . 
         [0024]    As shown in  FIG. 4 , the bifurcation  80  is a hollow duct-like structure through which various mechanical, electrical, pneumatic, hydraulic, or other connections (including structural supports) can pass without causing disruption to the airflow through the bypass duct  40 . In a typical installation of the gas turbine engine  10  under the wing of an aircraft (not shown), the upper bifurcation houses the engine mounts and various electrical, hydraulic, and pneumatic systems while the lower bifurcation houses oil drains and the like. The bifurcations “fair” or guide the flow in aerodynamic fashion around these structures. As will become apparent with respect to  FIGS. 5-7 , the forward edge of the bifurcation  80  is leaned and/or swept to meet with and blend into the trailing edge  73  of the guide vane  70 . The remaining portion of the bifurcation, aft of the leading edge portion, may be similarly leaned or may be more radially oriented as needed to accommodate structural loads and the passage of the various service connections. 
         [0025]      FIG. 5  shows the aerodynamic integration of the guide vane  70 , and particularly the trailing edge of the guide vane  70 , into the leading edge of the bifurcation  80 . This is an important aspect of implementing the swept and leaned guide vane designs into the integral vane frame engine architecture. 
         [0026]    Lean and/or sweep of the guide vanes and bifurcations may provide aerodynamic, acoustic, and/or other benefits in terms of gas turbine engine performance. Angles of sweep S such as about 0 to about 40 degrees aft, relative to the hub radial direction (normal to the central axis), and/or circumferentially leaning the outlet guide vane  70  with lean angles L from about −40 to about 0 degrees, relative to the radial orientation, in the direction of fan rotation, may provide acoustic benefits, such as reductions in noise from the fan assembly  12 . Angles of sweep greater than about 5 degrees aft, and angles of lean greater than about −5 degrees, are believed to be particularly useful. Negative angles of sweep, i.e., forward sweep, is also possible for some applications in comparable angular ranges of about −40 to about 0 degrees forward. Positive lean angles are also possible, in comparable angular ranges of about 0 to about 40 degrees. For the sake of illustration, the drawing figures depict configurations employing a lean angle of about −10 degrees and a sweep angle of about 25 degrees. 
         [0027]    Because of the sweep incorporated into the geometry of guide vane  70 , the axial location of the vane leading edge  71  varies with radial station. Aerodynamic and acoustic optimization also requires different vane turning angles at each radial station.  FIGS. 6 and 7  illustrate the extremes of these differences, at the vane root  72  and tip  74  locations, respectively. Comparison of these figures also illustrates increased axial fan/vane spacing at the vane tip  74 , due to the swept design in a positive (rearward) sweep configuration, which provides acoustic benefit.  FIGS. 4 and 7  depict the axial component of vane sweep, from root to tip, as the distance A. 
         [0028]      FIG. 8  is an elevational partial cross-sectional view similar to  FIG. 2  of another embodiment of the gas turbine engine assembly shown in  FIG. 1 . In  FIG. 8 , the same numbering scheme for individual elements described above with respect to  FIG. 2  is employed. The configuration of  FIG. 8  differs from that of  FIG. 2  in that the fan frame assembly  60  includes guide vanes  70  along with a smaller number (such as, for example, 6) structural strut members  90  spaced annularly around the bypass duct and a bifurcation  80 . In such a configuration, the strut members  90  are load-bearing elements which reduce the structural loads imparted to the guide vanes  70 . As previously described above, the strut members are incorporated into the bifurcation(s) and the guide vane(s) adjacent to the bifurcation(s) are blended or faired in such that the trailing edge of the guide vane and the leading edge of the respective bifurcation are blended together.  FIG. 9  is a perspective view of the gas turbine engine of  FIG. 8  in a typical installation configuration for an aircraft (not shown). 
         [0029]    The guide vanes and bifurcations may be fabricated from any suitable materials using any suitable fabrication methods as are known in the art and suitable for the intended configuration and operating environment. Configuration details, such as the number and thickness of guide vanes  70 , may influence the degree to which lean and sweep can be implemented without interfering with adjacent vanes. 
         [0030]    While much of the discussion has focused on an aviation gas turbine engine as the context for integration of the guide vane and bifurcation, it is foreseeable that such geometries and integrations may be suitable for use in other environments wherein a stationary guide vane and bifurcation are located downstream from rotating turbomachinery, such as wind or steam turbines. 
         [0031]    While the invention has been described in terms of various specific embodiments, those skilled in the art will recognize that the invention can be practiced with modification within the spirit and scope of the claims.

Technology Category: 2