Abstract:
In order to start a turbine engine ( 10 ), high-pressure fluid is directed onto a turbine ( 34   a ) to cause rotation of the turbine and thereby start the turbine engine. In a disclosed embodiment, the high-pressure fluid is provided through a fluid outlet ( 120 ) in a vane ( 36   a ) positioned adjacent the turbine ( 34   a ). The high-pressure fluid is provided by an air source, which may be another turbine engine, especially where the turbine engine to be started is a tip turbine engine that is not the primary propulsion source.

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 starter for a turbine engine, such as 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 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 pressure 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 pressure shaft. 
         [0004]    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. 
         [0005]    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 compressors. 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. 
         [0006]    Conventional turbine engines typically include a starter mounted to an external accessory drive gearbox. However, conventional starters cannot be used in the tip turbine engine because the tip turbine engine does not include an external accessory drive gearbox. 
       SUMMARY OF THE INVENTION 
       [0007]    In a turbine engine according to the present invention, high-pressure air is provided to the turbine to cause rotation of the turbine to start the turbine engine. The high-pressure air is provided through a fluid outlet upstream of the turbine. In the disclosed embodiment, the fluid outlet is provided in a vane positioned adjacent the turbine. 
         [0008]    The high-pressure air is provided by an air source, which may be a compressor, stored compressed air, an air starter, a gas generation device (pyrotechnic or other), or another turbine engine. In applications where the turbine engine to be started is a tip turbine engine that is not the primary propulsion source, the high-pressure air may be provided by the primary turbine engine. For example, the high-pressure air may be supplied from the bleed air from the primary 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 the engine centerline; and 
           [0012]      FIG. 3  is an enlarged view of the first stage of the turbine vanes and blades 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   a - b  (two stages shown) which rotatably drive the hollow fan blades  28  relative a plurality of tip turbine vanes  36   a - b  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  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 each 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 . 
         [0019]    The fan-turbine rotor assembly  24  includes a fan hub  64  that supports a plurality of the hollow fan blades  28 . Each hollow 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  in a direction 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 hollow fan blade section  72  where the airflow is centrifugally compressed. The diffuser section  74  receives the airflow from the core airflow passage  80 , and then diffuses the airflow and turns it 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. 
         [0020]    The tip turbine engine  10  further includes an air source  82  of high-pressure fluid, such as air, for starting the tip turbine engine  10 . The air source  82  is connected to a conduit  84  that provides the compressed air to a point just upstream of the first stage of the turbine blades  34   a.  In the embodiment shown, the conduit leads to the first stage of vanes  36   a,  just upstream from the turbine blades  34   a.  The air source  82  could be a compressor, stored compressed air, an air starter, a gas generation device (such as a pyrotechnic device or other), or other self-contained air source  82 . A valve  86  between the air source  82  and the first stage of vanes  36   a  allows for modulated fluid flow from the source  82  during the engine starting sequence. The valve  86  is closed once the engine  10  is started and is self-sustaining. In applications where the tip turbine engine  10  is not the only turbine engine, the air source  82  could be a conduit or reservoir connected to another turbine engine  210 . The other turbine engine  210  may be the primary means of propulsion, while the tip turbine engine  10  provides lift, control and/or supplementary propulsion. The other turbine engine  210  may be a conventional turbine engine or another tip turbine engine. In this case, high-pressure air for the air source  82  may be provided from the bleed air from the other turbine engine  210 . 
         [0021]    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  22  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  22 , and a planet carrier  94 , which rotates with the fan-turbine rotor assembly  24 . A plurality of first planet gears  93  each engage the sun gear  92  and a rotationally fixed ring gear  95 . The first 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. 
         [0022]      FIG. 3  illustrates in more detail the interaction of the vanes  36   a  and turbine blades  34   a  (one of each shown). The vane  36   a  is supplied with high-pressure air via the conduit  84 . The vane  36   a  includes a fluid outlet  120  at a trailing edge  122 , though flow could also exit from slots on the pressure (convex) and/or suction (concave) sides of the airfoil. When high-pressure air is supplied before the tip turbine engine  10  is started, the fluid outlet  120  directs the high-pressure air directly onto the turbine blade  34   a  in a generally axial direction, which urges the turbine blade  34   a  rotatably about the axis of the turbine  32  ( FIG. 2 ). 
         [0023]    Referring to  FIG. 2 , rotation of the turbine blades  34   a  causes rotation of the fan turbine rotor assembly  24 , which in turn causes the axial compressor rotor  46  to rotate. Rotation of the compressor blades  52  and the hollow fan blades  28  provides compressed air to the annular combustor  30 , which can then initiate ignition of the fuel and normal operation of the tip turbine engine  10 . 
         [0024]    When the combustor  30  ignition takes place and the engine  10  is self-sustaining, the valve  86  is closed. Once in normal operation, the core airflow enters the axial compressor  22 , where it is compressed by the compressor blades  52 . 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 . The diffuser section  74  receives air from the core airflow passage  80 , and turns and diffuses the airflow 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. 
         [0025]    The high-energy gas stream is expanded over the plurality of tip turbine blades  34   a - b  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 an exhaust case  106 . 
         [0026]    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   a - b  with the bypass airflow through the fan blades  28 . 
         [0027]    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.