Abstract:
A tip turbine engine ( 30 ) includes a low pressure compressor ( 22 ) having a plurality of inlet guide vanes ( 55 ) that are mounted at an inlet to the compressor case ( 50 ). Each inlet guide vane ( 55 ) includes at least one fluid outlet ( 56 ) proximate a trailing edge of the inlet guide vane ( 55 ), such that fluid flow through the fluid outlet ( 56 ) modulates and controls the air flow into the compressor ( 22 ). A supply of pressurized fluid may be supplied from compressed air from the compressor ( 22 ).

Description:
[0001]    This invention was conceived in performance of U.S. Air Force contract F33657-03-C-2044. The government may have rights in this invention. 
     
    
     BACKGROUND OF THE INVENTION 
       [0002]    The present invention relates to turbine engines, and more particularly to a jet flap inlet guide vane for a compressor for a tip turbine engine. 
         [0003]    An aircraft gas turbine engine of the conventional turbofan type generally includes a forward bypass fan, a low pressure compressor, a middle core engine, and an aft low pressure turbine, all located along a common longitudinal axis. A high pressure compressor and a high pressure turbine of the core engine are interconnected by a high spool shaft. The high pressure compressor is rotatably driven to compress air entering the core engine to a relatively high pressure. This high pressure air is then mixed with fuel in a combustor, where it is ignited to form a high energy gas stream. The gas stream flows axially aft to rotatably drive the high pressure turbine, which rotatably drives the high pressure compressor via the high spool shaft. The gas stream leaving the high pressure turbine is expanded through the low pressure turbine, which rotatably drives the bypass fan and low pressure compressor via a low pressure shaft. 
         [0004]    Some conventional gas turbine engines use mechanically activated, pivotably mounted inlet guide vanes at the compressor inlet to change the compressor airflow. However, these mechanically activated inlet guide vanes are heavy and costly. One conventiorial gas turbine engine includes a plurality of fixedly mounted inlet guide vanes, each including a plurality of holes adjacent a trailing edge. Compressed air taken from the compressor is fed to the inlet guide vanes and flows through the holes. The air through the holes in the inlet guide vanes redirects the inlet air flow without physically moving the inlet guide vanes. Controlling the amount of air supplied to the inlet guide vanes modulates and controls the inlet air flow. 
         [0005]    Although highly efficient, conventional gas turbine engines operate in an axial flow relationship. The axial flow relationship results in a relatively complicated elongated engine structure of considerable length relative to the engine diameter. This elongated shape may complicate or prevent packaging of the engine into particular applications. 
         [0006]    A recent development in gas turbine engines is the tip turbine engine. Tip turbine engines may include a low pressure axial compressor directing core airflow into hollow fan blades. The hollow fan blades operate as a centrifugal compressor when rotating. Compressed core airflow from the hollow fan blades is mixed with fuel in an annular combustor, where it is ignited to form a high energy gas stream which drives the turbine that is integrated onto the tips of the hollow bypass fan blades for rotation therewith as generally disclosed in U.S. Patent Application Publication Nos.: 20030192303; 20030192304; and 20040025490. The tip turbine engine provides a thrust-to-weight ratio equivalent to or greater than conventional turbofan engines of the same class, but within a package of significantly shorter length. 
       SUMMARY OF THE INVENTION 
       [0007]    A tip turbine engine includes a low pressure compressor having a plurality of inlet guide vanes that are mounted at an inlet to the compressor case. Each inlet guide vane includes at least one fluid outlet. Pressurized fluid through the at least one fluid outlet modulates and controls the flow of air into the compressor, without physically moving the inlet guide vanes. Thus, the inlet guide vanes are lighter weight and require fewer parts than the previously known methods. The supply of pressurized fluid may be supplied from compressed core air flow from the compressor. The low pressure compressor is mounted radially inward of the bypass air flow path. 
         [0008]    Because the compressor in a tip turbine engine is radially inward of a bypass air flow path, space in and around the compressor case is lrnited. The inlet guide vane of the present invention is simple, compact and lightweight and can be mounted within the compressor case of a tip turbine engine. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS  
         [0009]    Other advantages of the present invention can be understood by reference to the following detailed description when considered in connection with the accompanying drawings wherein: 
           [0010]      FIG. 1  is a partial sectional perspective view of a tip turbine engine. 
           [0011]      FIG. 2  is a longitudinal sectional view of the tip turbine engine of  FIG. 1  along an engine centerline. 
           [0012]      FIG. 3  is an enlarged top perspective sectional view of the compressor inlet guide vane of  FIG. 2 . 
       
    
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS  
       [0013]      FIG. 1  illustrates a general perspective partial sectional view of a tip turbine engine (TTE) type gas turbine engine  10 . The engine  10  includes an outer nacelle  12 , a rotationally fixed static outer support structure  14  and a rotationally fixed static inner support structure  16 . A plurality of fan inlet guide vanes  18  are mounted between the static outer support structure  14  and the static inner support structure  16 . Each inlet guide vane preferably includes a variable trailing edge  18 A. 
         [0014]    A nosecone  20  is preferably located along the engine centerline A to improve airflow into an axial compressor  22 , which is mounted about the engine centerline A behind the nosecone  20 . 
         [0015]    A fan-turbine rotor assembly  24  is mounted for rotation about the engine centerline A aft of the axial compressor  22 . The fan-turbine rotor assembly  24  includes a plurality of hollow fan blades  28  to provide internal, centrifugal compression of the compressed airflow from the axial compressor  22  for distribution to an annular combustor  30  located within the rotationally fixed static outer support structure  14 . 
         [0016]    A turbine  32  includes a plurality of tip turbine blades  34  (two stages shown) which rotatably drive the hollow fan blades  28  relative a plurality of tip turbine stators  36  which extend radially inwardly from the rotationally fixed static outer support structure  14 . The annular combustor  30  is disposed axially forward of the turbine  32  and communicates with the turbine  32 . 
         [0017]    Referring to  FIG. 2 , the rotationally fixed static inner support structure  16  includes a splitter  40 , a static inner support housing  42  and a static outer support housing  44  located coaxial to said engine centerline A. 
         [0018]    The axial compressor  22  includes the axial compressor rotor  46 , which is mounted for rotation upon the static inner support housing  42  through an aft bearing assembly  47  and a forward bearing assembly  48 . A plurality of compressor blades  52   a - c  extend radially outwardly from the axial compressor rotor  46 . A fixed compressor case  50  is mounted within the splitter  40 . A plurality of compressor vanes  54   a - c  extend radially inwardly from the compressor case  50  between stages of the compressor blades  52   a - c.  The compressor blades  52   a - c  and compressor vanes  54   a - c  are arranged circumferentially about the axial compressor rotor  46  in stages (three stages of compressor blades  52   a - c  and compressor vanes  54   a - c  are shown in this example). 
         [0019]    A plurality of compressor inlet guide vanes (IGVs)  55  are disposed upstream of the compressor blades  52   a - c  and compressor vanes  54   a - c.  A plurality of openings or nozzles  56  are formed near the trailing edge of the guide vanes  55 . The nozzles  56  are directed in a direction at approximately 45 degrees relative to the surface of the compressor IGV  55 . 
         [0020]    Some compressed air is supplied from the axial compressor  22  via conduit  58  to an optional jet valve  65 , which sends a controlled amount of the core air flow to the inlet guide vanes  55 . The jet valve  65  may adjust the amount of air flowing toward the inlet guide vanes  55  and may release excess air into the cavity between the compressor case  50  and the splitter  40 , where it may pass through the inlet guide vane  18  and discharge at an outer diameter of the nacelle  12 . 
         [0021]    The fan-turbine rotor assembly  24  includes a fan hub  64  that supports a plurality of the hollow fan blades  28 . Each fan blade  28  includes an inducer section  66 , a hollow fan blade section  72  and a diffuser section  74 . The inducer section  66  receives airflow from the axial compressor  22  generally parallel to the engine centerline A and turns the airflow from an axial airflow direction toward a radial airflow direction. The airflow is radially communicated through a core airflow passage  80  within the fan blade section  72  where the airflow is centrifugally compressed. From the core airflow passage  80 , the airflow is diffused and turned once again by the diffuser section  74  toward an axial airflow direction toward the annular combustor  30 . Preferably, the airflow is diffused axially forward in the engine  10 ; however, the airflow may alternatively be communicated in another direction. 
         [0022]    The tip turbine engine  10  may optionally include a gearbox assembly  90  aft of the fan-turbine rotor assembly  24 , such that the fan-turbine rotor assembly  24  rotatably drives the axial compressor rotor  46  via the gearbox assembly  90 . In the embodiment shown, the gearbox assembly  90  provides a speed increase at a 3.34-to-one ratio. The gearbox assembly  90  may be an epicyclic gearbox, such as a planetary gearbox as shown, that is mounted for rotation between the static inner support housing  42  and the static outer support housing  44 . The gearbox assembly  90  includes a sun gear  92 , which rotates the axial compressor rotor  46 , and a planet carrier  94 , which rotates with the fan-turbine rotor assembly  24 . A plurality of planet gears  93  each engages the sun gear  92  and a rotationally fixed ring gear  95 . The planet gears  93  are mounted to the planet carrier  94 . The gearbox assembly  90  is mounted for rotation between the sun gear  92  and the static outer support housing  44  through a gearbox forward bearing  96  and a gearbox rear bearing  98 . The gearbox assembly  90  may alternatively, or additionally, reverse the direction of rotation and/or may provide a decrease in rotation speed. 
         [0023]    A plurality of exit guide vanes  108  are located between the static outer support housing  44  and the rotationally fixed exhaust case  106  to guide the combined airflow out of the engine  10  and provide forward thrust. An exhaust mixer  110  mixes the airflow from the turbine blades  34  with the bypass airflow through the fan blades  28 . 
         [0024]      FIG. 3  illustrates one of the compressor IGVs  55  in more detail. The compressor IGV  55  includes an elongated interior chamber  111  in fluid communication with the nozzles  56 . Alternatively, conduit or other passageways could be defined within the compressor IGV  55 . Although the nozzles  56  are shown aligned proximate a trailing edge of the IGV  55 , other locations and configurations could be utilized. 
         [0025]    In operation, core airflow enters the axial compressor  22 , where it is compressed by the compressor blades  52 . As determined by the jet valve  65 , some of the core air flow is sent to the interior chambers  111  of the compressor IGVs  55 . This pressurized air then exits the nozzles  56  of the compressor IGVs  55 , thereby modulating and controlling the flow of air into the axial compressor  22 . The jet flap compressor IGVs  55  improve the stability of the tip turbine engine  10 , while providing a simply, lightweight, inexpensive means for providing such control. 
         [0026]    The compressed air from the axial compressor  22  that is not sent to the IGVs  55  enters the inducer section  66  in a direction generally parallel to the engine centerline A, and is then turned by the inducer section  66  radially outwardly through the core airflow passage  80  of the hollow fan blades  28 . The airflow is further compressed centrifugally in the hollow fan blades  28  by rotation of the hollow fan blades  28 . From the core airflow passage  80 , the airflow is turned and diffused axially forward in the engine  10  into the annular combustor  30 . The compressed core airflow from the hollow fan blades  28  is mixed with fuel in the annular combustor  30  and ignited to form a high-energy gas stream. 
         [0027]    The high-energy gas stream is expanded over the plurality of tip turbine blades  34  mounted about the outer periphery of the fan-turbine rotor assembly  24  to drive the fan-turbine rotor assembly  24 , which in turn rotatably drives the axial compressor  22  either directly or via the optional gearbox assembly  90 . The fan-turbine rotor assembly  24  discharges fan bypass air axially aft to merge with the core airflow from the turbine  32  in the exhaust case  106 . 
         [0028]    In accordance with the provisions of the patent statutes and jurisprudence, exemplary configurations described above are considered to represent a preferred embodiment of the invention. However, it should be noted that the invention can be practiced otherwise than as specifically illustrated and described without departing from its spirit or scope.