Abstract:
Improved annular components and improved methods for assembling annular components into a turbine engine are described with respect to an axial compressor having a plurality of annular compressor rotor airfoil assemblies ( 120 ) as an example. Each compressor rotor airfoil assembly comprises an annular rotor portion ( 122 ), a spacer portion ( 124 ) extending axially therefrom and a plurality of airfoils ( 52 ) extending radially therefrom. The plurality of airfoils may be integrally formed with the annular portion. The compressor rotor airfoil assemblies are stacked sequentially on a center-tie ( 134 ) or outer circumferential tie. The spacer portion of one compressor rotor airfoil assembly ( 120   a ) abuts the annular rotor portion of the adjacent compressor rotor airfoil assembly ( 120   b ) to retain one another on the center-tied outer circumferential tie. By stacking the compressor rotor airfoil assemblies sequentially and then retaining them, the typical split cases, flanges and rotor bolts may be eliminated.

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 improved annular components, such as axial compressor components for a turbine engine, and methods of assembling same in a 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 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. 
         [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 that receive core airflow 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. 
         [0006]    Both conventional and tip turbine engines may include a low pressure axial compressor. Such low pressure axial compressors include a plurality of axial compressor rotor blade assemblies each having a compressor rotor and a plurality of compressor blades extending radially therefrom. Conventionally, each blade is separately cast to include an airfoil portion and a root portion. The root portion of each conventional blade is slidably received within one of a plurality of grooves on the axial compressor rotor and is retained therein by an enlarged portion of the root portion. These conventional root connections increase the overall weight of the axial compressor rotor blade assemblies, as do the conventional connections between the multiple axial compressor rotor blade assemblies themselves. Therefore, lighter weight connections between the blades and the rotor in axial compressor rotor blade assemblies, and between the multiple axial compressor rotor blade assemblies themselves, would be desirable. 
       SUMMARY OF THE INVENTION 
       [0007]    This invention relates to improved annular components for turbine engines and improved methods for assembling such annular components into turbine engines. In one non-limiting embodiment, a turbine engine according to the present invention provides an improved compressor rotor blade assembly and an improved method for assembling compressor rotor blade assemblies into the axial compressor of a tip turbine engine. These compressor rotor blade assemblies each include an annular rotor portion and an integral spacer portion extending axially therefrom. A plurality of compressor blades extend radially from the annular rotor portion and are preferably machined from a single block of material or otherwise integrally formed with the rotor portion to form a continuous full hoop/ring component for each compressor stage. 
         [0008]    Each compressor rotor blade assembly is stacked sequentially on a rotor center-tie along the axis of the axial compressor. The spacer portion of each compressor rotor blade assembly abuts the rotor portion of the adjacent compressor rotor blade assembly to retain the adjacent rotor blade assembly on the rotor center-tie. By stacking the compressor rotor blade assemblies sequentially and then retaining them, the typical split cases and the rotor bolts can be (but need not be) eliminated. Eliminating split case flanges and bolts reduces the weight and cost of the turbine engine. Since all the split case flanges can be eliminated, this design also lends itself to counter-rotating axial compressor and/or turbine designs where split cases would have structural difficulties. 
     
    
     
       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 perspective view of a tip turbine engine, partially broken away. 
           [0011]      FIG. 2  is a partial longitudinal sectional view of the tip turbine engine of  FIG. 1  along the engine centerline A. 
           [0012]      FIG. 3  is a schematic front view of a portion of one of the compressor rotor blade assemblies of  FIG. 2 . 
           [0013]      FIG. 4  is a sectional view of the compressor rotor blade assembly of  FIG. 3  taken along line  4 - 4 . 
           [0014]      FIG. 5  is an enlarged sectional view of the axial compressor of  FIG. 2 . 
           [0015]      FIG. 6  is an enlarged, exploded sectional view of a portion of the axial compressor of  FIG. 5 . 
       
    
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS 
       [0016]      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.    
         [0017]    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 . 
         [0018]    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 . 
         [0019]    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. 
         [0020]    The axial compressor  22  includes the axial compressor rotor blade assembly  46  having a plurality of inner compressor blades  52  extending radially outwardly, and a fixed compressor case  50 . A plurality of outer compressor vanes  54  extend radially inwardly from the fixed compressor case  50  between stages of the inner compressor blades  52 . In this description and in the claims, blades, vanes or other airfoils in compressors or otherwise are referenced generically as “airfoils.” The inner compressor blades  52  and outer compressor vanes  54  are arranged circumferentially about the axial compressor rotor blade assembly  46  in stages (three stages of inner compressor blades  52  and three stages of outer compressor vanes  54  are shown in this example). The axial compressor rotor blade assembly  46  is mounted for rotation upon the static inner support housing  42  through a forward bearing assembly  68  and an aft bearing assembly  62 . 
         [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 . In operation, core airflow enters the axial compressor  22 , where it is compressed by the rotation of the inner 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 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 by rotation of the hollow fan blades  28 . 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. 
         [0022]    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. 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  via an optional gearbox assembly  90 . 
         [0023]    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 . 
         [0024]    The optional 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 . In the embodiment shown, the speed increase is at a 3.34-to-one ratio. The gearbox assembly  90  may be an epicyclic gearbox, such as a planetary gearbox as shown, that provides rotating engagement between the fan-turbine rotor assembly  24  and an axial compressor rotor blade assembly  46 . 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 the axial compressor  22 , and a planet carrier  94 , which rotates with the fan-turbine rotor assembly  24 . A plurality of planet gears  93  each engage 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 sun gear  92  is rotationally engaged with the axial compressor rotor blade assembly  46  at a splined interconnection  100  or the like. 
         [0025]    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 blade assembly  46 , or reverse rotational direction between the fan-turbine rotor assembly  24  and the axial compressor rotor blade assembly  46  via a plurality of second planet gears between the planet gears  93  and the ring gear  95 . 
         [0026]    As will be explained more fully below, the compressor rotor blade assembly  46  of the axial compressor  22  includes a plurality of compressor rotor blade assemblies  120 , one of which is shown in  FIGS. 3 and 4 . Each compressor rotor blade assembly  120  includes a plurality of inner compressor blades  52  integrally formed with an annular rotor portion  122 , such as by machining the inner compressor blades  52  and the rotor portion  122  from a single block of material. As can be seen more clearly in  FIG. 4 , an annular spacer portion  124  extends axially from the rotor portion  122  and has an inner radius r, that is greater than an inner radius r 2  of the rotor portion  122 , thereby defining a recess  130  radially inwardly of the spacer portion  124 . A pair of annular seals  128  may project radially outwardly from the spacer portion  124 . In the embodiment shown, the annular seals  128  are integrally-formed with the spacer portion  124  such that they rotate with the compressor blades  52  and seal against the inner diameter of the compressor vanes  54 . Because the bolted flanges have been eliminated, the torque required to drive the inner compressor blades  52  is now carried from one compressor rotor blade assembly  120  to the adjacent one, using either friction and/or some type of torque carrying feature machined into the rearward end  125  of the spacer portion  124  and/or the mating forward end  127  of the rotor portion  122 . One such feature is shown in  FIGS. 3 and 4  as a series of interlocking axial projections  126  disposed about the circumference of the rearward end  125  of the spacer portion  124 . Complementary interlocking recesses  132  could be disposed in the mating forward end  127  of the rotor portion  122  of the rearwardly adjacent compressor rotor blade assembly  120 . 
         [0027]    Referring to  FIGS. 5 and 6 , the axial compressor  22  includes a plurality of the compressor rotor blade assemblies  120   a - c , referenced as rear, middle and front compressor rotor blade assemblies  120   a - c , respectively, for clarity. The compressor rotor blade assemblies  120   a - c  are mounted on a generally conical rotor center-tie  134  or hub having inner and outer diameters that increase from an externally-threaded forward end  140  to a rearward end  142 . The outer surface  150  of the rotor center-tie  134  includes a plurality of cylindrical portions  144   a - c  that are generally parallel to the engine centerline A between conical portions  146   a - c . The rear compressor rotor blade assembly  120   a  has the largest inner radius r a  and the front compressor rotor blade assembly  120   c  has the smallest inner radius r c . The middle compressor rotor blade assembly  120   b  has an inner radius r b  sized between the other two. The rotor portion  122   a - c , particularly the inner surface  153   a - c  of the rotor portion  122   a - c , of each compressor rotor blade assembly  120   a - c  is generally parallel to the engine centerline A, although it should be understood that some slight taper might be helpful for assembly. The spacer portion  124   a - c , particularly the inner surface  154   a - c  of the spacer portion  124   a - c , is generally parallel to the conical portion  146   a - c  (i.e. parallel to the angle of the increase in diameter of the rotor center-tie  134   a - c ). 
         [0028]    Referring more specifically to  FIG. 6 , for assembly, the rear compressor rotor blade assembly  120   a  is first slid onto the rotor center-tie  134 , until the rotor portion  122   a  is mated with the cylindrical portion  144   a  of the rotor center-tie  134 . When mounted, the spacer portion  124   a  of the compressor rotor blade assembly  120   a  defines the recess  130   a  with the conical portion  146   a  of the rotor center-tie  134 . The middle compressor rotor blade assembly  120   b  is subsequently slid onto rotor center-tie  134  until the rotor portion  122   b  mates with the cylindrical portion  144   b , and the spacer portion  124   b  abuts the adjacent rotor portion  122   a  of the rear compressor rotor blade assembly  120   a . The front compressor rotor blade assembly  120   c  is then slid onto the rotor center-tie  134 , with the rotor portion  122   c  mounted on the cylindrical portion  144   c  and with the spacer portion  124   c  abutting the rotor portion  122   b  of the adjacent middle compressor rotor blade assembly  120   b.    
         [0029]    In this manner, compressor rotor blade assemblies  120   a - c  are stacked on the rotor center-tie  134  and retain one another on the rotor center-tie  134 . A nut  158  or other retaining device may be threaded or otherwise attached to an end, (e.g. the forward end  140 ) of the rotor center-tie  134 , thereby retaining all of the compressor rotor blade assemblies  120   a - c  on the rotor center-tie  134 . 
         [0030]    Depending upon the configuration of the outer compressor vanes  54 , the outer compressor vanes  54  may need to be assembled into the axial compressor in between mounting each of the compressor rotor blade assemblies  120   a - c . The outer compressor vanes  54  could be held together with bolted flanges, or the outer compressor vanes  54  could also use the stacked rotor assembly configuration illustrated and described with respect to the inner compressor blade assemblies  120   a - c . Although the compressor rotor blade assemblies  120   a - c  and center rotor-tie  134  are shown as used in a tip turbine engine  10 , they could also be used in a conventional turbine engine. Furthermore, while low pressure compressor rotor blade assemblies were described herein in detail, the stacking arrangement of this invention may also be used with low and/or high pressure compressor vane assemblies. Furthermore, these stacking arrangements may also be used in counter-rotating compressor and/or turbine designs. 
         [0031]    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.