Abstract:
A gas turbine engine ( 10 ), including: a turbine having radial inflow impellor blades ( 38 ); and an array of advanced transition combustor assemblies arranged circumferentially about the radial inflow impellor blades ( 38 ) and having inner surfaces ( 34 ) that are adjacent to combustion gases ( 40 ). The inner surfaces ( 34 ) of the array are configured to accelerate and orient, for delivery directly onto the radial inflow impellor blades ( 38 ), a plurality of discrete flows of the combustion gases ( 40 ). The array inner surfaces ( 34 ) define respective combustion gas flow axes ( 20 ). Each combustion gas flow axis ( 20 ) is straight from a point of ignition until no longer bound by the array inner surfaces ( 34 ), and each combustion gas flow axis ( 20 ) intersects a unique location on a circumference defined by a sweep of the radial inflow impellor blades ( 38 ).

Description:
STATEMENT REGARDING FEDERALLY SPONSORED DEVELOPMENT 
     Development for this invention was supported in part by Contract No. DE-FC26-05NT42644, awarded by the United States Department of Energy. Accordingly, the United States Government may have certain rights in this invention. 
    
    
     FIELD OF THE INVENTION 
     The invention relates to radial inflow gas turbine engines. In particular the invention relates to radial inflow gas turbine engines that utilize advanced transition combustion assemblies that do not utilize a first row of turbine vanes to accelerate combustion gases. 
     BACKGROUND OF THE INVENTION 
     Radial inflow gas turbine engines may have a turbine that uses an impeller to receive combustion gases from combustors and associated conventional transition ducts. The combustion gases rotate the impeller as the impeller directs the flow of combustion gases from a radially inward direction to an axial direction. Due to the nature of combustors and conventional transitions the combustion gases may be properly oriented by a first row of vanes disposed between an outlet of the conventional transition duct and the impeller. The first row of vanes may also accelerate the combustion gases to an appropriate speed. Such configurations with conventional transitions and a first row of turbine blades add cost, complexity, and reduce efficiency of the engine. Consequently, there remains room in the art for improvement. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The invention is explained in the following description in view of the drawings that show: 
         FIG. 1  shows a radial inflow gas turbine engine with advanced transition ducts. 
         FIG. 2  shows a close-up of the radial inflow gas turbine engine of  FIG. 1 . 
         FIG. 3  shows a cross section along line A-A of  FIG. 2 . 
         FIG. 4  shows a close-up of the circled region of  FIG. 3 . 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     The present inventor has recognized that the advanced transition duct concepts may be applied to a radial inflow gas turbine engine in order to provide a radial inflow engine that may be more compact (shorter), more efficient, and may cost less than conventional radial inflow engines. Advanced transition ducts applied to a radial inflow may be oriented radially, and within the radial plane may also be oriented such that the gas flow axis is also tangential to an impingement surface of the turbine blade at some point in the blade&#39;s rotation about a central axis of the gas turbine engine. Proper orientation of a flow of the combustion gases is a result of a geometry that is configured to create a straight axis for each combustion gas flow from a point of ignition until exiting the transition duct immediately prior to impinging the first row turbine blades. The combustion gas flow path is narrowed at an acceleration geometry portion along the straight axis. This narrowing accelerates the gas flow to a speed appropriate for delivery to the first row of turbine blades, without the need for any stationary vane in the flow path. A collimating geometry may be disposed between the accelerating geometry and an outlet of the advanced transition duct. The collimating geometry produces a uniform gas flow such that an entire volume of the gas flow is flowing parallel to the gas flow axis. As a result of this radial orientation, the engine may be shorter along its central axis. This allows the compressor to be closer to the turbine, which reduces rotor dynamics and shaft vibration problems incurred by relatively longer shafts of the prior art. A shorter shaft length, which allows the bearings to be closer together, is especially important with a radial inflow engine with a single impellor turbine. In such an engine the impellor is typically heavy. This necessitates a heavier shaft, which is even more susceptible to the rotor dynamics and vibration problems. Eliminating the first row of turbine vanes may also greatly reduce the cost of the turbine system of the engine, and efficiency losses associated with the first row vanes. 
       FIG. 1  depicts a radial inflow gas turbine engine  10  with a single impellor  12 , a rotor shaft  14 , a compressor  16  (an axial compressor is shown but a radial compressor is also envisioned), a combustor  17  and an associated advanced transition duct  18 . Each advanced transition duct  18  includes a gas flow axis  20  that is straight and disposed in a plane  22  that is perpendicular to a central axis  24  of the gas turbine engine  10 . Each advanced transition duct  18  may have an accelerating geometry  26  and a collimating geometry  28 . The accelerating geometry  26  may resemble a cone that decreases a size of cross sections of a flow path  32  in a downstream direction with respect to a direction of flow of the combustion gases within the advanced transition duct  18 . It is visible that the accelerating geometry narrows at least in a direction corresponding to the central axis  24 . 
     A plurality of advanced transition ducts  18  form an array (not shown), and inner surfaces  34  of the array form each flow path  32 . As used herein the term “inner” refers to surfaces in contact with combustion gases  40 . Each flow path  32  may be defined by inner surfaces  36  of a single advanced transition duct  18 , or a plurality of advanced transition ducts  18  working in conjunction. For example, a first advanced transition duct  18  may form part of a flow path  32 , and an upstream and or downstream adjacent (with respect to a direction of travel of the impellor  12  and impellor blades  38 ) advanced transition duct (not shown) may form a remainder of the flow path  32 . Combustion gases  40  flow along the gas flow axis  20 , are accelerated in the accelerating geometry  26  to a speed appropriate for deliver to the impellor blades  38 , and may be properly oriented in the collimating geometry  28 . Combustion gases  40  impinge the impellor blades  38 , and the impellor blades  38  in turn reorient the combustion gases  40  from a radial flow direction to an axial flow direction. 
     As shown in  FIG. 2 , each advanced transition duct  18  may be secured on either an advanced transition duct aft side  42  or an advanced transition duct fore side  44 . In the embodiment shown the advanced transition duct aft side  42  is secured to the engine turbine section casing  46 . Securing the advanced transition duct  18  on only one side or the other side minimizes or avoids thermal fright between forward and aft mounts. Also shown in greater detail are the accelerating geometry  26  and the collimating geometry  28 . 
       FIG. 3  depicts a partial cross section A-A of the engine  10  at the perpendicular plane  22  in which the gas flow axes  20  lie. A portion of the array  50  of advanced transition ducts  18  is visible. In this drawing a reference advanced transition duct  52  is circumferentially adjacent an upstream adjacent advanced transition duct  54 . The reference advanced transition duct  52  is also circumferentially adjacent a downstream adjacent advanced transition duct  56 . Upstream and downstream as used with adjacent advanced transition ducts is with respect to a direction of rotation  58  of the impellor  12  and impellor blades  38 . In the embodiment shown the reference advanced transition duct  52  comprises a reference duct inner surface  60 . The upstream adjacent advanced transition duct  54  has an upstream duct inner surface  62 , and the downstream adjacent advanced transition duct  56  has a downstream duct inner surface  64 . 
     Generally associated with each advanced transition duct  18  is a flow path  32 . However, each flow path  32  may be defined by inner surfaces of one or more than one advanced transition duct. In the embodiment shown, a reference flow path  66  is defined by the reference duct inner surface  60  as well as by the downstream duct inner surface  64 . Likewise, an upstream adjacent flow path  68  is defined by the upstream duct inner surface  62  and the reference duct inner surface  60 . The downstream adjacent flow path  70  is defined by the downstream duct inner surface  64  and an inner surface of an advanced transition duct disposed adjacent and downstream thereof. 
     It can be seen that in this embodiment the accelerating region  26  also narrows in a circumferential direction. Narrowing the both the axial and circumferential direction enables a significant acceleration of the combustion gases  40  and eliminates the need for the first row of turbine vanes. In each advanced transition duct  18  there is also a collimating geometry  28 . In this region the flow path  32  is fully bounded by the reference duct inner surface  60 , although more than once surface could define the collimating geometry  28 . Also in the depicted embodiment in each flow path  32  is a partially bounded region  72  where the combustion gases  40  are only partially bounded by the reference duct inner surface  60  and the downstream duct inner surface  64 . The collimating geometry  28  is upstream with respect to the flow of combustion gases  40  of the partially bounded region  72  because a partially bounded flow may diverge to a much greater degree if an entire volume of the flow is not flowing parallel to the gas flow axis  20 . However, since the entire volume of the flow is flowing parallel to the gas flow axis  20  after collimation, the flow will retain its cross sectional shape further downstream after exiting a flow path outlet  74  while traveling toward the impellor  12 . As a result of this tighter control of the combustion gases  40 , more the combustion gases  40  will impinge the impellor blades  38  in the manner desired to transfer the most energy, and thus the engine will operate more efficiently. 
     The flow path outlet  74  may be contoured circumferentially to match a profile made by a sweep  76  of a radially outward most point of the impellor blade tips  78 . The flow path outlet  74  may further be contoured along the central axis  24  to match an axial profile of the impellor blade tips  78 . 
     In the embodiment shown an upstream side  80  of the reference flow path  66  is adjacent a downstream side  82  of the upstream adjacent flow path  68 . These adjacent sides  80 ,  82  meet at a common geometry  84 . In this embodiment the common geometry  84  may form an edge that may be aerodynamically sharp. Aerodynamically sharp as used herein refers to a geometry that eliminates, or reduces to a negligible amount, a volume between the adjacent gas flow paths  66 ,  68 , where a wake of combustion gases  40  may form turbulence. An aerodynamically blunt geometry on the other hand may be a rounded corner, which would allow turbulence to occur between adjacent gas flow paths  66 ,  68 . Advantageously, an aerodynamically sharp geometry also enables a gap  86  to be minimized. Minimizing gap  86  directs the combustion gases  40  directly onto the impellor blades  38  almost immediately after the combustion gases  40  exit the flow path outlet  74 , reducing energy losses associate with longer travel distances. Further, since the reference flow path  66  and the upstream adjacent flow path  68  are at an angle with respect to each other, they converge on each other. Shortening the distance to the impellor blades  38  decreases the amount that the adjacent combustion gas flows converge and therefore interfere with each other, which again increases aerodynamic efficiency. 
       FIG. 4  is a close-up of the common geometry  84  that forms an edge. In this embodiment it is also in this region that adjacent advanced transition ducts  18  secure to each via interlocking geometry  88 . It is this securing that locks the individual adjacent advanced transition ducts  18  into the array  50 . The reference advanced transition duct  52  includes a hook geometry  90  and the upstream adjacent advanced transition duct  54  includes a hook receiving geometry  92 . The reference duct inner surface  60  meets the upstream duct inner surface  62  at an inner surface joint  94 . In this embodiment the inner surface joint  94  is disposed within the partially bounded region  72 , and consequently it is beneficial that the inner surface joint  94  be aerodynamically smooth, since a partially bound flow is more susceptible to perturbations. As used herein an aerodynamically smooth joint adds no flow disturbances (e.g. eddies and vortices), or a negligible amount, to the combustion gases  40  flowing over the inner surface joint  94 . Optimally the inner surface joint  94  would maintain a laminar flow within the combustion gases  40 . 
     The application of an advanced transition duct to a radial inflow gas turbine engine, and the novel and innovative structure required to accomplish this, enable advanced technology to be used within existing technology engine frames. As a result, more compact, simpler, more efficient, and less expensive radial inflow gas turbine engines may be achieved. Consequently, this represents an improvement in the art. 
     While various embodiments of the present invention have been shown and described herein, it will be obvious that such embodiments are provided by way of example only. Numerous variations, changes and substitutions may be made without departing from the invention herein. Accordingly, it is intended that the invention be limited only by the spirit and scope of the appended claims.