Abstract:
A tip turbine engine includes an axial compressor having a plurality of airfoils compressing core airflow. The airfoils include bleed air openings on their suction side surfaces. The bleed air openings prevent separation of the compressed airflow, which permits each airfoil stage to perform increased compression without separation of the airflow. As a result, the number of stages can be reduced, thereby shortening the overall length of the turbine engine.

Description:
BACKGROUND OF THE INVENTION 
       [0001]    The present invention relates to a turbine engine, and more particularly to an improved compressor for a tip turbine engine. 
         [0002]    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. 
         [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 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 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. 
         [0005]    Although much shorter axially than conventional turbine engines, much of the length of the tip turbine engine results from the number of stages in the axial compressor. Reducing the number of compressor stages would further decrease the axial length of the tip turbine engine. 
         [0006]    The number of stages could be reduced by using larger chord compressor blades that do more work in turning and compressing the air. However, at some point, the compressor blade tends to separate from the blade and the blade becomes highly inefficient, and can result in engine stall. 
         [0007]    Aspirated compressors have been used in conventional turbine engines to reduce the number of stages required in the compressor. In an aspirated compressor, suction is provided at selected locations on the surface of the compressor blades. The suction keeps the flow attached to the blade even with increased curvature and longer blade chord lengths. Aspirated compressors have not been implemented in tip turbine engines, which already have a shorter axial dimension. 
       SUMMARY OF THE INVENTION 
       [0008]    The present invention provides a tip turbine engine including an axial compressor having a plurality of airfoils compressing core airflow. The airfoils include bleed air openings on their suction side surfaces. The bleed air openings prevent separation of the compressed airflow, which permits each airfoil stage to perform increased compression without separation of the airflow. As a result, the number of stages can be reduced, thereby shortening the overall length of the turbine engine. 
         [0009]    In the example shown, the bleed air openings of the compressor blades are connected to a low pressure area radially outward of the combustor, which also provides a cool layer of air between the combustor and the adjacent airframe structure. The bleed air openings of the compressor vanes are connected to a low pressure area in an air-oil heat exchanger for cooling lubrication for a gearbox in the turbine engine. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0010]    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: 
           [0011]      FIG. 1  is a partial sectional perspective view of a tip turbine engine according to the present invention. 
           [0012]      FIG. 2  is a partial longitudinal sectional view of the tip turbine engine of  FIG. 1  taken along an engine centerline. 
       
    
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS 
       [0013]      FIG. 1  illustrates a general perspective partial sectional view of a tip turbine engine 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 vanes  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 , 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  aft of the compressor blades  52 . 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 . Although in the embodiment shown only a single stage of compressor blades  52  and a single stage of compressor vanes  54  are necessary, a plurality of stages of compressor blades  52  and compressor vanes  54  may be provided; however, overall, the number of stages of compressor blades  52  and/or compressor vanes  54  can be reduced with the present invention. 
         [0019]    The compressor blades  52  and the compressor vanes  54  are larger and provide more turning than previous designs, such that sufficient compression is provided in the single stage. In order to prevent separation, each of the compressor blades  52  and each of the compressor vanes  54  include at least one bleed opening  55  on its suction surface  56 . The bleed opening  55  may be a slot, as shown, or a plurality of holes. The bleed opening  55  on the compressor blade  52  leads through the interior of the compressor blade  52  to an aperture  57  at the tip of the compressor blade  52 . The tip of the compressor blade  52  is positioned adjacent an annular bleed chamber  58 . One or more conduits  59  lead from the annular bleed chamber  58  to a low-pressure area, which in the example shown is the region between the hot combustion chamber  30  and the adjacent airframe structure. One or more conduits  60  lead from the bleed openings  55  on the compressor vanes  54  to another low-pressure area, which in the example shown is an air-oil heat exchanger  88  for cooling the lubrication system of the gearbox  90 . 
         [0020]    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  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 by the diffuser section  74  toward the annular combustor  30 . 
         [0021]    Generally, 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 tip 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 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 . 
         [0022]    A plurality of fuel injectors  82 , or “nozzles,” (one shown) supply fuel to the combustor  30 . Fuel is delivered to the fuel injectors  82  from a fuel manifold or ring  84  extending circumferentially about the engine centerline A. 
         [0023]    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 . 
         [0024]    In operation, referring to  FIG. 2 , air enters the axial compressor  22 , where it is compressed by the compressor blades  52  and compressor vanes  54 . Suction from the low-pressure areas is provided through the bleed openings  55  on the suction side surfaces  56  of the compressor blades  52  and the compressor vanes  54  via the conduits  59 ,  60 . The suction provided on the suction side surfaces  56  prevents a separation of the airflow from the airfoils (compressor blades  52  and compressor vanes  54 ) that would otherwise occur due to the large amount of turning and compression provided by the compressor blades  52  and compressor vanes  54 . 
         [0025]    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 the annular combustor  30 . The 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 bulkhead  118  to the combustion chamber  112 . The fuel is injected into the annular combustor  30  where it is mixed with the core airflow and ignited to form a high-energy gas stream. 
         [0026]    The high-energy gas stream expands through the turbine vanes  36  and the tip turbine blades  34 . 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 . 
         [0027]    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 tip turbine blades  34  with the bypass airflow through the fan blades  28 . 
         [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. For example, although the invention is shown as used in a tip turbine engine, the present invention would be beneficial in most or all conventional gas turbine engines.