Abstract:
A tip turbine engine ( 10 ) and a method of operating the engine provides increased efficiency while eliminating or educing the number of axial compressor ( 122 ) stages by moving the core airflow inlet ( 27 ) aft of the fan ( 24 ). As a result, the core airflow entering the core airflow inlet is the fan exhaust, which is already at an increased pressure. A portion of the fan exhaust is guided radially inward, then axially forward and then radially outward through compressor chambers ( 72 ) in the hollow fan blades ( 28 ) for further, centrifugal compression.

Description:
BACKGROUND OF THE INVENTION 
       [0001]    The present invention relates to a tip turbine engine, and more particularly to a tip turbine engine with a core airflow inlet aft of a bypass fan. 
         [0002]    An aircraft gas turbine engine of the conventional turbofan type generally includes a forward bypass fan, a low pressure compressor, a high pressure compressor, a combustor, a high pressure turbine, and a low pressure turbine, all located along a common longitudinal axis. The low and high pressure compressors are rotatably driven to compress entering air to a relatively high pressure. This high pressure air is then mixed with fuel in the combustor, where it is ignited to form a high energy gas stream. This gas stream flows axially aft to rotatably drive the high pressure turbine, which rotatably drives the high pressure compressor via a high spool shaft. The gas stream leaving the high pressure turbine is expanded through the low pressure turbine, which rotatably drives the forward bypass fan and the low pressure compressor via a low spool shaft. 
         [0003]    Although highly efficient, conventional turbofan engines operate in an axial flow relationship. The axial flow relationship results in a relatively complicated elongated engine structure of considerable longitudinal length relative to the engine diameter. This elongated shape may complicate or prevent packaging of the engine into particular applications. 
         [0004]    A recent development in gas turbine engines is the tip turbine engine. Tip turbine engines include hollow fan blades through which core airflow flows, such that the hollow fan blades operate as centrifugal compressor chambers. 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 at least equivalent to conventional turbofan engines of the same class, but within a package of significantly shorter length. 
         [0005]    In some tip turbine engine designs, core airflow may be compressed by an axial compressor before entering the hollow fan blades for further, centrifugal compression. The axial compressor may include an axial compressor rotor with one or more stages of radially-extending compressor blades. Increasing the number of stages of compressor blades increases the compression of the core airflow and the efficiency of the engine, but increases the overall length and weight of the engine and the number of parts. 
       SUMMARY OF THE INVENTION 
       [0006]    A tip turbine engine according to the present invention provides increased efficiency while eliminating or reducing the number of axial compressor stages by moving the core airflow inlet aft of the fan. As a result, the core airflow entering the core airflow inlet is the fan exhaust, which is already compressed by the fan. The fan exhaust is fed axially forward and then radially outward through compressor chambers in the hollow fan blades for further, centrifugal compression. 
         [0007]    The tip turbine engine may optionally include an axial compressor between the core airflow inlet and the compressor chambers in the hollow fan blades. However, in contrast to axial compressors that are located forward of a bypass fan, these axial compressors utilize high-pressure fan exhaust, which allows them to have fewer stages therein while still providing the same high pressure core airflow to the compressor chambers in the hollow fan blades. 
         [0008]    In one embodiment, the fan of the tip turbine engine drives the axial compressor via at least one gear that increases the rate of rotation of the axial compressor relative to the fan and/or reverses the direction of rotation of the axial compressor relative to the fan. 
     
    
     
       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 longitudinal sectional view of a first embodiment of a tip turbine engine according to the present invention. 
           [0011]      FIG. 2  is a longitudinal sectional view of a second embodiment of a tip turbine engine according to the present invention. 
           [0012]      FIG. 3  schematically illustrates an alternate gearbox assembly that could be used in the tip turbine engine of  FIG. 2 . 
       
    
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS 
       [0013]      FIG. 1  illustrates a partial sectional view of a tip turbine engine (TTE) type gas turbine engine  10  taken along an engine centerline A. 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 fan-turbine rotor assembly  24  is mounted for rotation about the engine centerline A fore of a core airflow passage  26  having a core airflow inlet  27 . The fan-turbine rotor assembly  24  includes a plurality of hollow fan blades  28  to provide internal, centrifugal compression of the compressed airflow for distribution to an annular combustor  30  located within the rotationally fixed static outer support structure  14 . The core airflow inlet  27  is aft of the fan blades  28  and leads to the core airflow passage  26 , which reverses the core airflow such that it flows back toward the fan-turbine rotor assembly  24  in a direction generally parallel to the engine centerline A. 
         [0015]    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 . 
         [0016]    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 traveling generally parallel to the engine centerline A from the core airflow passage  26 , 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 hollow fan blade section  72 , which acts as a compressor chamber where the airflow is centrifugally compressed. From the core airflow passage  80 , the airflow is diffused and turned once again 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. 
         [0017]    In operation, airflow enters the engine  10  and passes between inlet guide vanes  18  and rotating fan blades  28 . The rotating fan blades  28  compress the airflow and discharge high-pressure fan exhaust. A portion of the fan exhaust enters the core airflow inlet  27  and is reversed by the core airflow passage  26 . The core airflow passage  26  turns the axially rearward flowing fan exhaust radially inwardly and then axially forward toward the inducer section  66 . The reversed core airflow 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 , where it is ignited to form a high-energy gas stream. 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 . Concurrent therewith, the fan-turbine rotor assembly  24  discharges fan bypass air (fan exhaust) axially aft to merge with the core airflow from the turbine  32  in an exhaust case  106 . A plurality of exit guide vanes  108  extend inwardly from the rotationally fixed static outer support structure  14  to guide the combined airflow out of the engine  10  and provide forward thrust. An exhaust mixer  109  mixes the airflow from the turbine blades  34  with the bypass airflow through the fan blades  28 . 
         [0018]    By feeding back some of the high-pressure fan exhaust as the core airflow, the efficiency of the engine  10  is increased, without the need for an axial compressor. This reduces the overall length and weight of the engine  10  and reduces the number of parts. 
         [0019]      FIG. 2  illustrates a second embodiment of a tip turbine engine  110  according to the present invention which additionally incorporates an axial compressor  122  for even further compression of the core airflow. Components that are similar to those described above with respect to  FIG. 1  are indicated with the same reference numeral, and the description of those components and their operation is incorporated by reference here. 
         [0020]    The axial compressor  122  is mounted between the core airflow passage  26  and the inducer sections  66 . The axial compressor  122  includes an axial compressor rotor  146 , from which a plurality of compressor blades  152  extend radially outwardly, and a fixed compressor case  150 . A plurality of compressor vanes  154  extend radially inwardly from the compressor case  150  between stages of the compressor blades  152 . The compressor blades  152  and compressor vanes  154  are arranged circumferentially about the axial compressor rotor  146  in stages (two stages of compressor blades  152  and compressor vanes  154  are shown in this example). 
         [0021]    The axial compressor rotor  146  may be driven by the fan-turbine rotor assembly  24  either directly, or via a gearbox assembly  190 , as shown. The gearbox assembly  190  shown provides a speed increase between the fan-turbine rotor assembly  24  and the axial compressor  122 , at a ratio of 3.34 to 1, for example. The gearbox assembly  190  may include a planetary gearset, including a sun gear  192  coupled to the axial compressor rotor  146  and a planet carrier  194  coupled to the fan-turbine rotor assembly  24  to provide a speed differential therebetween. A plurality of planet gears  193  (one shown) are mounted to the planet carrier  194 . The planet gears  193  engage the sun gear  192  and a ring gear  195 . Rotating the axial compressor rotor  146  at a rate higher than that of the fan-turbine rotor assembly  24  increases the compression provided by the axial compressor  122 . The gearbox assembly  190  could alternatively provide a speed decrease between the fan-turbine rotor assembly  24  and the axial compressor rotor  146 . 
         [0022]    An alternative gearbox assembly  290  that reverses the direction of rotation between the fan-turbine rotor assembly  24  and the axial compressor  122  is shown schematically in  FIG. 3 . The gearbox assembly  290  provides second planet gears  198  coupled between each planet gear  193  and the ring gear  195  and mounted to the planet carrier  194 . The gearbox assembly  290  is otherwise similar to gearbox assembly  190  as described above. The gearbox assembly  290  may also provide a speed increase or a speed decrease. 
         [0023]    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.