Abstract:
A tip turbine engine has a more efficient core airflow path from the hollow fan blades, through the combustor and to the combustion chamber of a combustor. The turbine engine includes a rotatable fan having a plurality of radially-extending fan blades each defining compressor chambers extending radially therein. A turbine is mounted to the outer periphery of the fan. A diffuser at a radially outer end of each compressor chamber turns core airflow through the compressor chamber toward the combustor. The high velocity, high pressure core airflow from the compressor chambers in the hollow fan blades is diffused before the compressed core airflow enters the combustor, thereby improving the efficiency of the tip turbine engine. Further, the overall diameter of the tip turbine engine is reduced by the arrangement of the diffuser case in a position not directly radially outward of the fan blades.

Description:
The present invention relates to a tip turbine engine, and more particularly to a novel core airflow path for a tip turbine engine. 
     An aircraft gas turbine engine of the conventional turbofan type generally includes a forward bypass fan and 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 spool shaft. 
     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 length relative to the engine diameter. This elongated shape may complicate or prevent packaging of the engine into particular applications. 
     A recent development in gas turbine engines is the tip turbine engine. Tip turbine engines locate an axial compressor forward of a bypass fan which includes hollow fan blades that receive airflow from the axial compressor therethrough such that the hollow fan blades operate as a centrifugal compressor. 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. 
     In the known tip turbine engines, most of the core airflow flows radially outwardly from the radial outer ends of the hollow fan blades into the combustor, which is mounted about the periphery of the fan. A fuel injector aft of the fan delivers fuel into the combustor where it is ignited. The high-energy gas stream is then directed axially forward in the combustor, then redirected radially inward and then turned once again axially rearward to pass through turbine blades between the fan blades to rotatably drive the fan. This design has some drawbacks. Mounting the combustor about the periphery of the fan increases the overall diameter of the known tip turbine engine. 
     Additionally, in the known tip turbine engines, the compressed airflow from the hollow fan blades exits directly into the combustor. A lack of diffusion between the centrifugal compressor and the combustor causes a large loss in efficiency. 
     SUMMARY OF THE INVENTION 
     A tip turbine engine according to the present invention has a more efficient core airflow path from the hollow fan blades through to the combustion chamber of the combustor. 
     The turbine engine includes a rotatable fan having a plurality of radially-extending fan blades each defining compressor chambers extending radially therein. A turbine is mounted to the outer periphery of the fan. A diffuser at a radially outer end of each compressor chamber turns core airflow through the compressor chamber toward an annular combustor disposed in front of the turbine. Thus, in the tip turbine engine according to the present invention, the high velocity, high pressure core airflow from the compressor chambers in the hollow fan blades is diffused before the compressed core airflow enters the combustor, thereby improving the efficiency of the tip turbine engine. Further, the overall diameter of the tip turbine engine is reduced by the arrangement of the diffuser case in a position not directly radially outward of the fan blades. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       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: 
         FIG. 1  is a partial sectional perspective view of a tip turbine engine according to the present invention. 
         FIG. 2  is a longitudinal sectional view of the tip turbine engine of  FIG. 1  taken along an engine centerline. 
         FIG. 3  is an enlarged view of the diffuser, combustor and turbine area of  FIG. 2 . 
     
    
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS 
       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. 
     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 . 
     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 . 
     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 . 
     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. 
     The axial compressor  22  includes the axial compressor rotor  46 , from which a plurality of compressor blades  52  extend radially outwardly, and a fixed compressor case  50 . A plurality of compressor vanes  54  extend radially inwardly from the compressor case  50  between stages of the compressor blades  52 . The compressor blades  52  and compressor vanes  54  are arranged circumferentially about the axial compressor rotor  46  in stages (three stages of compressor blades  52  and compressor vanes  54  are shown in this example). The axial compressor rotor  46  is mounted for rotation upon the static inner support housing  42  through a forward bearing assembly  68  and an aft bearing assembly  62 . 
     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  which acts as a compressor chamber 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. 
     All or substantially all of the airflow through the core airflow passage  80  is core airflow directed by the diffuser section  74  axially forward toward the combustor  30 . Minimal amounts of airflow may be directed radially outwardly from the diffuser section  74  through the turbine blades  34  (paths not shown) to cool the tip turbine blades  34 . This cooling airflow is then discharged through radially outer ends of the tip turbine blades  34  and then into the annular combustor  30 . However, at least substantially all of the airflow is core airflow directed by the diffuser section  74  toward the combustor  30 . As used herein, “core airflow” is airflow that flows to the combustor  30 . 
     A gearbox assembly  90  aft of the fan-turbine rotor assembly  24  provides a speed increase between the fan-turbine rotor assembly  24  and the axial compressor  22 , which in the embodiment shown is at a 3.34 ratio. In the embodiment shown, the gearbox assembly  90  is a planetary gearbox that provides co-rotating engagement between the fan-turbine rotor assembly  24  and an axial compressor rotor  46 . Alternatively, a counter-rotating planetary gearbox could be provided. The gearbox assembly  90  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 with the axial compressor  22 , and a planet carrier  94 , which rotates with the fan-turbine rotor assembly  24  to provide a speed differential therebetween. A plurality of planet gears  93  (one shown) are mounted to the planet carrier  94 . The planet gears  93  engage the sun gear  92  and a ring gear  95 . The gearbox assembly  90  is mounted for rotation between the sun gear  92  and the static outer support housing  44  through a forward bearing  96  and a rear bearing  99 . The forward bearing  96  and the rear bearing  99  are both tapered roller bearings and both handle radial loads. The forward bearing  96  handles the aft axial load, while the rear bearing  99  handles the forward axial loads. The sun gear  92  is rotationally engaged with the axial compressor rotor  46  at a splined interconnection  100  or the like. 
     It should be noted that the gearbox assembly  90  could utilize other types of gear arrangements or other gear ratios and that the gearbox assembly  90  could be located at locations other than aft of the axial compressor  22 . For example, the gearbox assembly  90  could be located at the front end of the axial compressor  22 . Alternatively, the gearbox assembly  90  could provide a speed decrease between the fan-turbine rotor assembly  24  and the axial compressor rotor  46 . 
     The annular combustor  30  and turbine  32  are shown in greater detail in  FIG. 3 . The annular combustor  30  is located entirely fore of a fan plane P, within which the fan blades  28  rotate. The annular combustor  30  includes an annular combustion chamber  112  defined between an annular inner combustion chamber wall  114  and annular outer combustion chamber wall  116 . A forward wall  118  at a forward end of the combustion chamber  112  has mounted thereto a fuel injector  120 , which directs fuel into the combustion chamber  112 . The combustion chamber  112  includes a combustion chamber outlet  122  opposite the forward wall  118 . The combustion chamber outlet  122  is substantially axially aligned with the forward wall  118  such that a substantially axial combustion path  124  is defined through the combustion chamber  112 . The annular inner and outer combustion chamber walls  114 ,  116  and the forward wall  118  are perforated to permit core airflow into the combustion chamber  112 . 
     An annular diffuser case  128  substantially encloses the annular inner and outer combustion chamber walls  114 ,  116  and the forward wall  118 . An inner diffuser case wall  130  defines a core airflow path  132  with the annular inner combustion chamber wall  114 . A core airflow path inlet  134  is axially aligned (i.e. along an axis parallel to the engine centerline A ( FIG. 1 )) with the diffuser section  74  and is substantially radially aligned (i.e. along a radius from engine centerline A) with the combustion chamber outlet  122 . The core airflow path inlet  134  leads into the combustion chamber  112  through the annular inner and outer combustion chamber walls  114 ,  116  and the forward wall  118 . 
     In operation, referring to  FIG. 2 , air enters the axial compressor  22 , where it is compressed by the three stages of the compressor blades  52  and compressor vanes  54 . The compressed air from the axial compressor  22  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  by diffuser section  74  into core airflow path inlet  134  of the annular combustor  30 , as shown in  FIG. 3 . This diffusion improves the efficiency, by reducing the losses encountered when the compressed core airflow enters the larger combustion chamber  112 . The diffused compressed core airflow from the hollow fan blades  28  then flows radially outwardly and through the annular inner and outer combustion chamber walls  114 ,  116  and the forward wall  118  to the combustion chamber  112  where it is mixed with fuel and ignited to form a high-energy gas stream. 
     The high-energy gas stream expands and follows the combustion path  124 , which is substantially axial all the way from the forward wall  118  of the combustion chamber  112  through the combustion chamber outlet  122  and through the tip turbine blades  36 . The high-energy gas stream rotatably drives 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 drives the axial compressor  22  via the gearbox assembly  90 . Because the combustion path  124  is substantially axial, the efficiency of the combustor  30  is improved over the known combustors in tip turbine engines. Additionally, because the combustor  30  is located fore of the fan blades  28  and is not located the fan plane P, the tip turbine engine  10  has a smaller diameter than the known tip turbine engines. 
     The fan-turbine rotor assembly  24  discharges fan bypass air axially aft to merge with the core airflow from the turbine  32  in an exhaust case  106 . A plurality of exit guide vanes  108  are located between the static outer support housing  44  and 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  110  mixes the airflow from the turbine blades  34  with the bypass airflow through the fan blades  28 . 
     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.