Patent Publication Number: US-10760483-B2

Title: Tip turbine engine composite tailcone

Description:
This application is a continuation of U.S. patent application Ser. No. 11/720,463 filed May 30, 2007 which was a National Application of PCT/US2004/040176 filed Jan. 12, 2004. 
    
    
     This invention was made with government support under Contract No. F33657-03-C-2044. The government therefore has certain rights in this invention. 
    
    
     BACKGROUND OF THE INVENTION 
     The present invention relates to a tip turbine engine, and more particularly to a non-metallic composite tailcone for the 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. Air flows into the engine at an inner diameter and an outer diameter of the engine. Air entering through the outer diameter of the engine exits the rear of the engine at the outer diameter as a relatively low temperature gas stream after flowing through the bypass fan. 
     Airflow entering the inner diameter of the engine enters a high pressure compressor driven by a high pressure turbine to compress the air to a high pressure. This high pressure air is then mixed with fuel in a combustor and ignited to form a relatively high temperature gas stream. The high temperature gas stream flows axially aft to rotatably drive the high pressure turbine which rotatably drives the high pressure compressor through the high pressure shaft. The high temperature gas stream leaving the high pressure turbine is expanded through the low pressure turbine which rotatably drives the bypass fan and low pressure compressor through a low pressure shaft. The high temperature gas stream exits the rear of the engine at the inner diameter from the low pressure turbine and flows over a tailcone. The tailcone is made of titanium or other metal to withstand the elevated temperatures of the high temperature gas stream. 
     Although highly efficient, conventional turbofan engines operate in an axial flow relationship where airflow essentially travels linearly through the engine. 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 in 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 that is radially outward from the fan blades and near an outside diameter of the tip turbine engine. The airflow and fuel mixture is ignited to form a relatively high temperature gas stream which drives the turbine integrated onto the tips of the hollow bypass fan blades for rotation therewith as disclosed in U.S. Patent Application Publication Nos. 2003192303; 20030192304; and 20040025490. Unlike conventional engines, the relatively high temperature gas stream exits the rear of the tip turbine engine at the outside diameter, while the relatively low temperature gas stream flowing through the bypass fan exits the rear of the tip turbine engine at the inner diameter and flows over the tailcone. 
     Accordingly and because of the low temperature gas stream discharged from the inner diameter in a tip turbine engine, it is desirable to provide a lighter, non-metallic tailcone in the tip turbine engine. 
     SUMMARY OF THE INVENTION 
     The non-metallic tailcone according to the present invention includes a wall structure disposed about a central axis. The wall structure defines an interior compartment and a forward portion that tapers to an aft portion of the non-metallic tailcone. The wall structure of the non-metallic tailcone is a polymer composite that includes woven Kevlar™ fiber impregnated with bismaleimide. 
     The tip turbine engine produces a first temperature gas stream from a first output source and a second temperature gas stream from a second output source. The first output source is a combustor and the second output source is a bypass fan. The second temperature gas stream from the bypass fan is a lower temperature than the first temperature gas stream from the combustor. The first temperature gas stream is discharged from the combustor at an outer diameter of the tip turbine engine. The second temperature gas stream is discharged from the bypass fan at an inner diameter of the tip turbine engine and flows over the tailcone mounted to a structural frame of the engine. Discharging the warmer first temperature gas stream at the outer diameter and the cooler second temperature gas stream at the inner diameter allows a non-metallic to be used to form the tailcone. 
     The present invention therefore takes advantage of the cooler temperature gas stream discharged from the inner diameter in a tip turbine engine by providing a light weight non-metallic tailcone in the tip turbine engine. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The various features and advantages of this invention will become apparent to those skilled in the art from the following detailed description of the currently preferred embodiment. The drawings that accompany the detailed description can be briefly described as follows: 
         FIG. 1  is a partial sectional perspective view of an exemplary tip turbine engine incorporating a non-metallic tailcone according to the present invention; 
         FIG. 2  is a cross-sectional view of the tip turbine engine of  FIG. 1 ; and 
         FIG. 3  is a cross-sectional view of a second embodiment of a non-metallic tailcone according to the present invention. 
     
    
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT 
       FIG. 1  illustrates a partial sectional perspective view of a tip turbine engine (TTE) type gas turbine engine  8 . The engine  8  includes an outer case  10 , a nonrotatable static outer support structure  12 , a nonrotatable static inner support structure  14 , and a non-metallic tailcone  16 . A plurality of fan inlet guide vanes  18  are mounted between the static outer support structure  12  and the static inner support structure  14 . 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 . The axial compressor  22  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 nonrotatable static outer support structure  12 . 
     A turbine  32  includes a plurality of tip turbine blades  34  (two stages shown) which rotatably drive the hollow fan blades  28  relative to a plurality of tip turbine stators  36  which extend radially inwardly from the nonrotatable static outer support structure  12 . The annular combustor  30  is axially forward of the turbine  32  and communicates with the turbine  32 . 
     Referring to  FIG. 2 , the nonrotatable static inner support structure  14  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 static outer support housing  44  includes exit guide vanes  46  that extend radially outward to the outer case  10 . 
     The axial compressor  22  includes the axial compressor rotor  48  from which a plurality of compressor blades  52  extend radially outwardly and a compressor case  50  fixedly mounted to the splitter  40 . 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  48  in stages (three stages of compressor blades  52  and compressor vanes  54  are shown in this example). The axial compressor rotor  48  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 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  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  where the airflow is centrifugally compressed. 
     From the core airflow passage  80 , the airflow is turned and diffused toward an axial airflow direction toward the annular combustor  30 . Preferably the airflow is diffused axially forward in the engine  8 , however, the airflow may alternatively be communicated in another direction. An exhaust mixer  82  extends from an exhaust case portion  84  of the outer case  10  and guides airflow exiting the annular combustor  30  over the tip turbine blades  34  and tip turbine stators  36 . 
     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 . 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 shaft  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. The gearbox assembly  90  is preferably a planetary gearbox that provides co-rotating or counter-rotating rotational engagement between the fan-turbine rotor assembly  24  and an axial compressor rotor  48 . The gearbox assembly  90  is mounted for rotation between the sun gear shaft  92  and the static outer support housing  44  through a forward bearings  96  and a rear bearing  98 . The forward bearings  96  and the rear bearing  98  are both tapered roller bearings and both handle radial loads. The forward bearing  96  handles the aft axial load, while the rear bearing  98  handles the forward axial loads. 
     The sun gear shaft  92  is rotationally engaged with the axial compressor rotor  48  at a splined interconnection  100  or the like. Alternatively, the gearbox assembly  90  could provide a speed decrease between the fan-turbine rotor assembly  24  and the axial compressor rotor  48 . 
     The non-metallic tailcone  16  is part of a tailcone assembly  102  that attaches to the static outer support housing  44  with a set of fasteners  104 , although only one fastener is illustrated in the  FIG. 2 . The tailcone assembly  102  houses a device  106 , such as an oil cooler or other device, and includes a frustoconical surface  108 . A wall structure  110  disposed about central axis  112  forms the frustoconical surface  108 . The wall structure  110  defines an interior compartment  114  and a forward portion  116  that tapers to an aft portion  118  of the tailcone assembly  102 . 
     A polymer forms the wall structure  110 . Preferably the polymer is a polymer composite and includes woven Kevlar™ fiber impregnated with bismaleimide. Other polymer composites such those utilizing epoxy, other polymers, carbon fibers, glass fibers, and other reinforcing substances may also be utilized. Use of a non-metal may provide the benefit of a lighter overall engine weight compared to conventional titanium or other metal tailcones, as the density of non-metals is generally lower than the density of metal. Alternatively, a ceramic matrix composite may be utilized, including ceramic matrix composites with metallic reinforcing substances. 
     In operation, air enters the engine  8  at 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 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  8  into a first output, an 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 first temperature gas stream. The first temperature 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 drives the axial compressor  22  through the gearbox assembly  90 . The first temperature gas stream then exits the tip turbine blades  34  and is guided by the exhaust mixer  82 . 
     Concurrent therewith, a second output, the fan-turbine rotor assembly  24 , discharges a second temperature gas stream at an inner engine diameter  128  to merge with the first temperature gas stream from the turbine  32 . The combined airflow exits the engine  8  over the frustoconical surface  108  of the non-metallic tailcone  16  to provide forward thrust. The cooler second temperature gas stream from the hollow fan blades  28  discharging at the inner engine diameter  128  over the tailcone assembly  102  allows the non-metallic tailcone  16  to be formed from polymer rather than more temperature resistant titanium or metal material as in conventional engines. 
       FIG. 3  shows a partial cross-sectional view of a second embodiment of a non-metallic tailcone according to the present invention. The non-metallic tailcone  202  is attached to the static outer support housing  44  with a set of fasteners  204 , although only one fastener is illustrated in the  FIG. 3 . The non-metallic tailc one  202  includes a frustoconical surface  208  and a wall structure  210  disposed about a central axis  212 . The wall structure  210  defines an interior compartment  214  and a forward portion  216  with a first radius R 1  that tapers to an aft portion  218  with a radius R 2  which is less than the diameter R 1 . Preferably, the non-metallic tailcone  202  includes woven Kevlar™ fiber impregnated with bismaleimide. Other composites such as those utilizing epoxy, other polymers, carbon fibers, glass fibers, other reinforcing substances, and ceramic matrix composites may also be utilized. Use of a non-metallic composite may provide the benefit of a lighter overall engine weight compared to conventional titanium or other metal tailcones, as the density of non-metals is generally lower than the density of metal. 
     It should be understood that relative positional terms such as “forward,” “aft,” “upper,” “lower,” “above,” “below,” and the like are with reference to the normal operational attitude of the vehicle and should not be considered otherwise limiting. 
     It should be understood that although a particular component arrangement is disclosed in the illustrated embodiment, other arrangements will benefit from the instant invention. 
     Although a preferred embodiment of this invention has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this invention. For that reason, the following claims should be studied to determine the true scope and content of this invention.