Patent Publication Number: US-2015086396-A1

Title: Turbocharger with mixed flow turbine stage

Description:
TECHNICAL FIELD 
     The present disclosure is directed to a turbocharger and, more particularly, to a turbocharger with a mixed flow turbine stage. 
     BACKGROUND 
     Internal combustion engines such as, for example, diesel engines, gasoline engines, and gaseous fuel powered engines are supplied with a mixture of air and fuel for subsequent combustion within the engines that generates a mechanical power output. In order to increase the power output generated by this combustion process, an engine can be equipped with a turbocharged air induction system. 
     A turbocharged air induction system includes a turbocharger that uses exhaust from the engine to compress air flowing into the engine, thereby forcing more air into a combustion chamber of the engine than the engine could otherwise draw into the combustion chamber. This increased supply of air allows for increased fueling, resulting in an increased power output. A turbocharged engine typically produces more power than the same engine without turbocharging. 
     A conventional turbocharger includes a turbine housing, and a turbine wheel centrally disposed within the housing and driven by exhaust to rotate a connected compressor wheel. The exhaust is pushed against blades connected to the turbine wheel to cause rotation of the turbine wheel. In some applications, vanes disposed on a nozzle ring connected to the turbine wheel accelerate the exhaust through the blades. The vanes and/or the blades of the turbine can direct the exhaust in axial, radial, and tangential directions. 
     A mixed flow turbine is generally viewed as across design between a radial and an axial turbine. An exemplary mixed flow turbine is disclosed in U.S. Pat. No. 8,128,356 to Higashimori that issued on Mar. 6, 2012 (the &#39;356 patent). Specifically, the &#39;356 patent describes a mixed flow turbine having blades whose outline of leading edges located at an upstream side is formed in a convex shape toward the upstream side, and a scroll that is a space formed upstream of the blades by a casing having a shroud that covers the radially external edges of the blades. Working fluid is supplied at a hub and the shroud and flows substantially in axial, radial, and tangential directions at a shroud-side inlet channel and at a hub-side inlet channel. A shape of the leading edges of the blades is designed to reduce incidence loss. 
     Although the mixed flow turbine of the &#39;356 patent may be adequate for some applications, it may still be less than optimal at wide operating conditions. In particular, the mixed flow turbine of the &#39;356 patent directs a non-uniform and poorly guided mixed flow through the turbine stage at wide operating conditions, which can result in high energy losses, reduced aerodynamic efficiencies, and increased mechanical or vibrational stresses (or strains) on the turbine during operation due to flow misalignment (high incidence) with the blades of the turbine. Also, the blade angle and thickness distributions of the mixed flow turbine shown in &#39;356 patent are generally not smooth like a Bezier curve, which can lead to problems manufacturing the blades. 
     The turbocharger of the present disclosure solves one or more of the problems set forth above and/or other problems of the prior art. 
     SUMMARY 
     In one aspect, the present disclosure is directed to a turbocharger. The turbocharger may include a housing at least partially defining a compressor shroud and a turbine shroud. The turbocharger may also include a compressor wheel disposed within the compressor shroud, a shaft connected to the compressor wheel, and a turbine wheel disposed within the turbine shroud and connected to an end of the shaft opposite the compressor wheel. The turbine wheel may have a generally annular hub, and a plurality of blades disposed radially around the hub. Each of the plurality of blades may include an airfoil having a hub face connected to the hub, a shroud face opposite the hub face and oriented towards the turbine shroud, a trailing edge, and a leading edge opposite the trailing edge. An angle between a base of the hub and the leading edge may be about 25-55 degrees. The leading edge may be substantially straight or substantially concave in a meridional plane. The turbocharger may also include a nozzle ring having a ring-shaped generally flat plate located at a periphery of the turbine wheel, and a plurality of vanes disposed radially around an upper surface of the plate. A camber of each of the plurality of vanes may be generally S-shaped along its meridional length from a leading edge to a trailing edge of each of the plurality of vanes. 
     In a second aspect, the present disclosure is directed to a turbine blade for a turbocharger. The turbine blade may include an airfoil having a hub face connected to a turbine wheel hub of the turbocharger, a shroud face located opposite the hub face and oriented towards a turbine shroud of the turbocharger, a trailing edge, and a leading edge opposite the trailing edge. An angle between a base of the turbine wheel and the leading edge may be about 25-55 degrees. The leading edge may be substantially straight or substantially concave in a meridional plane. 
     In a third aspect, the present disclosure is directed to nozzle ring for a turbocharger. The nozzle ring may include a ring-shaped generally flat plate having an inner annular hub, and an outer annular flange radially spaced apart from the inner annular hub. The nozzle ring may also include a plurality of vanes disposed between the inner annular hub and the outer annular flange. A camber of each of the plurality of vanes may be generally S-shaped along its meridional length from a leading edge to a trailing edge of each of the plurality of vanes. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a schematic illustration of an exemplary disclosed power system; 
         FIG. 2  is a cross-sectional illustration of an exemplary disclosed turbocharger that may be used in conjunction with the power system of Ea.  1 ; 
         FIG. 3  is a pictorial illustration of an exemplary disclosed turbine wheel and nozzle ring that may be used in conjunction with the turbocharger of  FIG. 2 ; 
         FIG. 4  is a side-view illustration of the turbine wheel of  FIG. 3 ; 
         FIG. 5  is a meridional-view illustration of an exemplary disclosed turbine blade that may be used in conjunction with the turbine wheel of  FIG. 4 ; 
         FIG. 6  is a meridional-view illustration of an alternative embodiment of the turbine blade of  FIG. 5 ; 
         FIGS. 7 and 8  are charts associated with exemplary disclosed geometry of the turbine blade of  FIGS. 5 ; and 
         FIGS. 9 ,  10 , and  11  are charts associated with exemplary disclosed geometry of the nozzle ring of  FIG. 3 . 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  illustrates a power system  10  having an engine  12 , an air induction system  14 , and an exhaust system  16 . For the purposes of this disclosure, engine  12  is depicted and described as a four-stroke diesel engine. One skilled in the art will recognize, however, that engine  12  may be any other type of combustion engine such as, for example, a two- or four-stroke gasoline or gaseous fuel-powered engine. Air induction system  14  may be configured to direct air or a mixture of air and fuel into engine  12  for combustion. Exhaust system  16  may be configured to direct combustion exhaust from engine  12  to the atmosphere. 
     Engine  12  may include an engine block  18  that at least partially defines a plurality of cylinders  20 . A piston (not shown) may be slidably disposed within each cylinder  20  to reciprocate between a top-dead-center position and a bottom-dead-center position, and a cylinder head (not shown) may be associated with each cylinder  20 . Each cylinder  20 , piston, and cylinder head may together at least partially define a combustion chamber. In the illustrated embodiment, engine  12  includes twelve cylinders  20  arranged in a V-configuration (i.e., a configuration having first and second banks  22 ,  24  or rows of cylinders  20 ). However, it is contemplated that engine  12  may include a greater or lesser number of cylinders  20  and that cylinders  20  may be arranged in an inline configuration, in an opposing-piston configuration, or in another configuration, as desired. 
     Air induction system  14  may include, among other things, at least one compressor  28  that may embody a fixed geometry compressor, a variable geometry compressor, or any other type of compressor configured to receive air and compress the air to a desired pressure level. Compressor  28  may direct air to one or more intake manifolds  30  associated with engine  12 . It should be noted that air induction system  14  may include multiple compressors  28  arranged in a serial configuration, a parallel configuration, or a combination serial/parallel configuration. 
     Exhaust system  16  may include, among other things, an exhaust manifold  34  connected to one or both of banks  22 ,  24  of cylinders  20 . Exhaust system  16  may also include at least one turbine  32  driven by the exhaust from exhaust manifold  34  to rotate compressor  28  of air induction system  14 . Compressor  28  and turbine  32  may together form a turbocharger  36 . Turbine  32  may be configured to receive exhaust and convert potential energy in the exhaust to a mechanical rotation. After exiting turbine  32 , the exhaust may be discharged to the atmosphere through an aftertreatment system  38  that may include, for example, a hydrocarbon closer, a diesel oxidation catalyst (DOC), a diesel particulate filter (DPF), and/or any other treatment device known in the art, if desired. It should be noted that exhaust system  16  may include multiple turbines  32  arranged in a serial configuration, a parallel configuration, or a combination serial/parallel configuration, as desired. 
     As illustrated in  FIG. 2 , compressor  28  and turbine  32  of turbocharger  36  may be connected to each other via a common shaft  50 . Turbocharger  36  may include a housing  40  at least partially defining compressor and turbine shrouds  42 ,  44  that are configured to house corresponding compressor and turbine wheels  46 ,  48 . Compressor shroud  42  may include an axially-oriented inlet  52  located at a first axial end  54  of turbocharger  36 , and a tangentially-oriented volute  56  located between first axial end  54  and a second axial end  58  of turbocharger  36 . Turbine shroud  44  may include a volute  60  located between volute  56  and second axial end  58  of turbocharger  36 . Turbine shroud  44  may be configured to receive exhaust flow from exhaust manifold  34  in a tangential direction at a volute inlet (not shown). Volute  60  may direct the exhaust flow in three directions: axially (along rotation axis X), radially inward (along a radius of the volute), and tangentially (around a rotation axis X) toward and through a nozzle ring  62 . Nozzle ring  62  may be disposed downstream of volute  60  and be configured to accelerate exhaust gas flowing therethrough. 
     As compressor wheel  46  is rotated, air may be drawn axially into turbocharger  36  via inlet  52  and directed toward compressor wheel  46 . Blades  64  of compressor wheel  46  may then push the air radially outward in a spiraling fashion and into intake manifolds  30  (referring to  FIG. 1 ) via an outlet volute (not shown). Similarly, as exhaust from exhaust system  16  is directed axially, radially, and tangentially inward toward turbine wheel  48 , the exhaust may push against blades  66  of turbine wheel  48 , causing turbine wheel  48  to rotate and drive compressor wheel  46  via shaft  50 . After passing through turbine wheel  48 , the exhaust flow may exit axially outward through a turbine outlet  68  located at second axial end  58  of turbocharger  36  into aftertreatment system  38  (shown only in  FIG. 1 ). 
     As illustrated in  FIG. 3 , turbine wheel  48  may be generally disc-shaped and include a generally annular hub  70 . Blades  66  may extend outward in three dimensions from annular hub  70 . Nozzle ring  62  may be located radially upstream of turbine wheel  48  (i.e., at a periphery of turbine wheel  48 ). While turbine wheel  48  rotates in a rotational direction R, nozzle ring  62  may be stationary. Nozzle ring  62  may be generally ring-shaped, and include an inner annular hub  72  and an outer annular flange  74 . A plurality of three-dimensional vanes  76  may be disposed between inner annular hub  72  and outer annular flange  74  to direct and accelerate exhaust flow from volute  60  toward blades  66  of turbine wheel  48 . 
     As shown in  FIG. 3 , each blade  66  may include an airfoil  78  having a lower face (also known as a hub face)  80  that is connected to hub  70 , an opposing upper face (also known as a shroud face)  82  that is oriented towards an inner surface of shroud  44 , a trailing edge  84  that is proximate to turbine outlet  68 , a leading edge  86  that is opposite to trailing edge  84 , a high-pressure side (also known as the pressure side)  88 , and an opposing low-pressure side (also known as the suction side)  90 . It is contemplated that trailing edge  84  may be located closer to turbine outlet  68  than leading edge  86 . 
     Similarly, each vane  76  may include a lower face (also known as a hub face)  92  that is connected to nozzle ring  62 , an opposing upper face (also known as a shroud face)  94  that is oriented towards an inner surface of shroud  44 , a trailing edge  96  located proximate to turbine wheel  48 , a leading edge  98  that is opposite to trailing edge  96 , a high-pressure side (also known as the pressure side)  100 , and an opposing low-pressure side (also known as the suction side)  102 . It is contemplated that trailing edge  96  may be located closer to turbine wheel  48  than leading edge  98 , 
       FIG. 4  illustrates a side-view of turbine wheel  48 . For the purposes of this disclosure, a blade forward sweep angle α B  of blade  66  may refer to an angle between leading edge  86  of blade  66  and a base of hub  70 . A meridional length L MB  of blade  66  may refer to a meridional distance between trailing and leading edges  84 ,  86  of blades  66  along a camber line passing through a lengthwise center of the blades. Blades  66  may curve along their lengths, each forming a corresponding meridional blade angle β B  defined by the following equation: 
     
       
         
           
             
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     θ=Angular coordinate, polar angle, or wrap angle 
     z m =Local meridional coordinate along the meridional length 
     r=Local radial location 
     β B =Local meridional blade angle 
     A thickness T B  may refer to a distance between low- and high-pressure sides  88 ,  90  that is generally orthogonal to the camber line. A spacing S B  may refer to a straight line distance between adjacent trailing edges  84  of adjacent blades  66 . A solidity ratio SR B  of blade  66  may be defined as the ratio of the meridional chord length L MB  to the spacing S B  (SR B =L MB /S B ). 
       FIG. 5  illustrates a meridional view of a single blade  66  taken along the meridional length L MB . In the meridional plane shown in  FIG. 5 , an R-axis defines a radial direction, and a Z-axis defines an axial direction along the meridional length.  FIG. 5  shows an inlet flow passage  104  adjacent to leading edge  86  (i.e., where exhaust enters leading edge  86  of blade  66 ), and an outlet flow passage or diffuser  106  adjacent to trailing edge  84  (i.e., where exhaust exits trailing edge  84  of blade  66 ).  FIG. 5  also shows a hub curve  108  corresponding to hub face  80  and a shroud curve  110  corresponding to shroud face  82 , it should be noted that the relationship between hub face  80  and shroud face  82  possess unique geometric and “ruled element blade” characteristics. In a ruled element blade, an angular location is defined by a straight line drawn in 3D space between points at span locations along the hub and shroud faces  80 ,  82 , it should also be noted that hub curve  108  and shroud curve  110  are master curves and control generation of all other defining curves (e.g., intermediate curves between hub curve  108  and shroud curve  110 ). Modification of the hub and/or shroud curves  108 ,  110  may result in a subsequent modification of the intermediate curves. 
     For the purposes of this disclosure, a blade inlet cone angle λ B  may refer to an angle between the R-axis of the meridional plane and leading edge  86  of blade  66 . An inlet hub radius r 4H  may refer to a distance from the Z-axis of the meridional plane to a point on the hub curve  108  at leading edge  86 . An inlet shroud radius r 4S  may refer to a distance from the Z-axis of the meridional plane to a point on the shroud curve  110  at leading edge  86 . An inlet width W B  may refer to a distance between the point on the hub curve  108  at leading edge  86  and the point on the shroud curve  110  at leading edge  86 . An inlet width ratio WR B  may be defined as the ratio of the width W B  to the meridional length L MB  (WR=W B /L MB ) AZ-axis offset Z B  may refer to a distance between the R-axis and the point on the hub curve  108  at leading edge  86 . A non-dimensional Z-axis offset ratio ZR may be defined as the ratio of the Z-axis offset Z B  to the meridional length L MB  (ZR B =Z B /L MB ). An exit deviation angle (or clip angle) δ B  may refer to an angle between trailing edge  84  of blade  66  and the R-axis of the meridional plane. An exit hub radius r 5H  may refer to a distance from the Z-axis of the meridional plane to a point on the hub curve  108  at trailing edge  84 . An exit shroud radius r SS  may refer to a distance from the Z-axis of the meridional plane to a point on the shroud curve  110  at trailing edge  84 . A turbine trim TR B  may be defined by the following equation: [(r 5X /r 4S ) 2 ×100)]. A diffuser hub exit radius r 6R  TR B  may refer to a distance from the Z-axis of the meridional plane to a point on the hub curve  108  at the diffuser  106 . A diffuser shroud exit radius r 5s  may refer to a distance from the Z-axis of the meridional plane to a point on the shroud curve  110  at the diffuser  106 . 
     The aerodynamic performance of a radial and mixed flow turbine is usually interpreted as a function of velocity ratio U/C 0 , where U is the blade tip speed and C 0  is the isentropic velocity, resulting from ideal expansion of gas through a pressure ratio equal to that of the turbine. Since turbochargers often need to operate at low U/C 0  operating conditions (or high expansion ratio conditions at constant tip speed), there is a need for an efficient turbine stage design to operate at these low U/C 0  conditions with low aerodynamic losses (e.g., incidence loss). The disclosed geometry of blade  66  has been selected to provide a desired aerodynamic flow uniformity and guidance through turbine  32  that reduces flow misalignment (incidence) and results in improved performance and efficiency at wide operating conditions (especially at low U/C 0  conditions) of turbocharger  36 . In addition, the disclosed geometry of blade  66  increases structural integrity and manufacturability of the blades. For example, each blade  66  may have a blade forward sweep angle α B  of about 25-55°. In one embodiment, the blade forward sweep angle α B  is about 47°. Blade  66  may also have a blade inlet cone angle λ B  of about 50-70°. In one embodiment, the blade inlet cone angle λ B  is about 58°. Blade  66  may further have a clip angle δ B  of about 0-14°. In one embodiment, the clip angle δ B  is about 7°. These angle ranges may help to reduce the incidence of exhaust flowing through turbine  32  and improve vibration characteristics of the turbine  32 , thereby improving aerodynamic performance and structural integrity of turbocharger  36 . 
     In the disclosed embodiment, the solidity ratio SR B  of blade  66  may be about 0.8-1.2, with about 10 to 17 blades  66  for a given turbine  32 . In one embodiment, the solidity ratio SR B  is about 1.05 for a turbine  32  housing  13  blades. The turbine trim TR B  of blade  66  may be about 50-80. In one embodiment, the turbine trim TR B  is about 59. The width ratio WR B  of blade  66  may be about 0.2-0.42. In one embodiment, the width ratio WR B  is about 0.29. The Z-axis offset ratio ZR B , of blade  66  may be about 0.07-0.20. In one embodiment, the Z-axis offset ratio ZR B  is about 0.13. Each of these geometrical features may help to improve aerodynamic performance and structural integrity of blades  66 , while at the same time allow for smooth curves that are conducive to improving manufacturability. In particular, these geometrical features may create a blade profile that is suitable for flank milling. 
     As described above,  FIG. 5  shows an inlet flow passage  104  adjacent to leading edge  86  (i.e., where exhaust enters leading edge  86  of blade  66 ), and an outlet flow passage or diffuser  106  adjacent to trailing edge  84  (i.e., where exhaust exits trailing edge  84  of blade  66 ). The disclosed geometry of inlet flow passage  104  has been selected to provide a desired aerodynamic flow guidance and uniformity into the turbine blade leading edge  86  that also reduces flow misalignment (incidence) and results in improved performance and efficiency at wide operating conditions (especially at low INC, conditions) of turbocharger  36 . The outlet flow passage  106  which acts like a diffuser  106  may have a diffuser ratio at hub (r 5h /r 6h ) from 1.15 to 1.55. In one embodiment, diffuser ratio at hub is 1.35. The outlet flow passage  106  which acts like a diffuser  106  may have a diffuser ratio at shroud (r 6s /r 5s ) from 1.02 to 1.10. In one embodiment, diffuser ratio at shroud is 1.07. 
       FIG. 6  shows an alternative embodiment of blade  66 . In this embodiment, leading edge  86  of blade  66  is substantially concave rather than being substantially straight as shown in the embodiment of  FIG. 5 . It is contemplated that having a concave leading edge  86  may help to further improve flow alignment at wide operating conditions, in some applications. 
     In order to further improve manufacturability and aerodynamic performance of blades  66 , the meridional blade angle β B  may change along the meridional length L MB . Specifically,  FIG. 7  shows a plurality of curves corresponding to the meridional blade angle β 3  between hub curve  108  and shroud curve  110 . It should be noted that each curve between hub curve  108  and shroud curve  110  may correspond with an intermediate layer of blade  66  between the hub and shroud faces  80 ,  82 . As can be seen from a comparison of the plurality of curves, the meridional blade angle β 11  at hub face  80  may be generally larger than the meridional blade angle β B  at shroud face  82  (i.e., blade  66  may be more vertical at hub face  80 ). In addition, the meridional blade angle β B  at both faces reaches a maximum at leading edge  86  and a minimum at a trailing edge  84 . In other words, the meridional blade angle β B  generally decreases from leading edge  86  to trailing edge  84 . As also shown in  FIG. 7 , the meridional blade angle β B  may vary between about −5° and 30° at leading edge  86 , and vary between about −40° and −80° at trailing edge  84 . This blade angle distribution may help to reduce aerodynamic losses and, thus, improve performance and efficiency of turbocharger  36 . 
     As shown in  FIG. 8 , the thickness T B  of blades  66  may also vary along their meridional length L MB . In particular,  FIG. 8  shows a plurality of curves corresponding to the thickness T B  of blades  66  between the hub and shroud faces  80 ,  82  relative to the meridional length L MB  of blades  66 . As can be seen from the plurality of curves, the thickness of blades  66  may reach a maximum thickness T Bmax  of about 10 mm at about 60-80% of the meridional length  140  and be thinnest at trailing and leading edges  84 ,  86 . In one embodiment, the maximum thickness T Bmax  may be at about 68% of the meridional length L MB . Also shown in  FIG. 8 , the thickness of blades  66  may be substantially greater along hub face  80  than along shroud face  82 . Finally, a maximum thickness at the leading edge  86  may be about 0.38×T Bmax , while a maximum thickness at the trailing edge  84  may be about 0.61×T Bmax . This smooth Bezier curve thickness distribution of blades  66  may improve the manufacturability of the blades, especially using flank milling processes, which can be lower in cost than alternative manufacturing processes. 
     Referring back to  FIG. 3 , exemplary disclosed geometry of vanes  76  of nozzle ring  62  will now be discussed. For the purposes of this disclosure, a meridional length L MV  may refer to a meridional distance between trailing and leading edges  96 ,  98  of vanes  76  along a camber line passing through a lengthwise center of the vanes. Similar to blades  66 , vanes  76  may also curve along their lengths, each forming a corresponding meridional vane angle β V  defined by the following equation: 
     
       
         
           
             
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     θ=Angular coordinate, polar angle, or wrap angle 
     z m =Local meridional coordinate along the meridional length 
     r=Local radial location 
     β V =Local meridional vane angle 
     A thickness T V  may refer to a distance between high- and low-pressure sides  100 ,  102  that is generally orthogonal to the camber line of vane  76 . A chord length L CV  may refer to a straight line distance between trailing and leading edges  96 ,  98  of vanes  76 . A spacing S V  may refer to a straight line distance between adjacent trailing edges  96  of adjacent vanes  76 . A solidity ratio SR V  may be defined as the ratio of the chord length L CV  to the spacing S V  (SR V =L CV /S V ). A width W V  may refer to a distance between hub face  92  and shroud face  94  at leading edge  86 . A width ratio WR V  may be defined as the ratio of the width W V  to the chord length L CV  (WR V =W V /L CV ). A blade inlet shroud tip radius r 1  may refer to a distance from a center of turbine wheel  48  to leading edge  86  of blade  66  at shroud face  82 . A vane leading edge radius r 2  may refer to a distance from the center of turbine wheel  48  to leading edge  98  of vane  76  at shroud face  82 . A vane inlet radius ratio My may be defined as the ratio of the vane leading edge radius r 2  to the blade inlet shroud tip radius r 1 . A nozzle inlet stagger angle φ V  may refer to an angle between the chord length L cv  and the vane leading edge radius r 2 . A vane trailing edge radius r 3  may refer to a distance from the center of turbine wheel  48  to trailing edge  96  of vane  76  at shroud face  82 . A vane exit radius ratio ER V  may be defined as the ratio of the vane trailing edge radius r 3  to the blade inlet shroud tip radius r 1 . 
     Similar to blades  66 , the disclosed geometry of vanes  76  has been selected to provide desired aerodynamic flow angles with improved flow uniformity at an exit of nozzle ring  62 , increased structural integrity of the vanes, and low torque loading of the vanes  76 . For example, each vane  76  may have a solidity ratio SR V  of about 0.7-1.2, with about 13 to 25 vanes  76  included around nozzle ring  62 . In one embodiment, the solidity ratio is about 1.11, with 23 blades included around nozzle ring  62 . The width ratio W RV  of vane  76  may be about 0.2-0.40. In one embodiment, the width ratio WR V  is about 0.23. The vane inlet radius ratio IR of vane  76  may be about 1.3-1.5. In one embodiment, the vane inlet radius ratio IR is about 1.36. The vane exit radius ratio ER of vane  76  may be about 1.05-1.3. In one embodiment, the vane inlet radius ratio ER is about 1.19. Finally, the nozzle inlet stagger angle φ V  of vane  76  may be about 60°-80°. In one embodiment, the nozzle inlet stagger angle φ V  is about 74°. Each of these geometrical features may help to reduce aerodynamic losses, reduce vane torque loading, and improve the structural integrity of vanes  76 , while at the same time allow for smooth curves that are conducive to improving manufacturability. 
     Also, similar to blades  66 , the meridional vane angle β V  of vanes  76  may change along the meridional length L MV . Specifically,  FIGS. 9 and 10  show curves corresponding to the meridional vane angle β V  from leading edge  98  to trailing edge  96  for two different embodiments of nozzle ring  62 . In each of these embodiments, the meridional vane angle β V  may vary in a range of about 50-80° along its meridional length. 
     In a first embodiment shown in  FIG. 9 , the meridional vane angle β V  may be substantially different along the hub and shroud faces  92 ,  94 . Thus, two separate curves are shown. A hub curve  112  may correspond to the meridional vane angle β V  along the hub face  92 , and a shroud curve  114  may correspond to the meridional vane angle β V  along the shroud face  94 . In this embodiment, both curves  112 ,  114  may share a substantially S-shaped curve along the meridional length, representing a shape of the chamber of vane  76 . However, the hub curve  112  may be substantially greater than the shroud curve  114  at each point along the meridional length between leading edge  98  and trailing edge  96 . 
     In a second embodiment shown in  FIG. 10 , the meridional vane angle γ V  may be substantially equal along the hub and shroud faces  92 ,  94 . Thus, only one curve is shown.  FIG. 10  shows a curve  116  that corresponds to the meridional vane angle β V  for both the hub and shroud faces  92 ,  94 . In this embodiment, vane  76  may also have a generally S-shaped camber. Further, in this embodiment, there may be an inclination angle from leading edge  98  to trailing edge  96 , shown here as inclination curve  118 . Both of the above embodiments of vanes  76  may be used with nozzle ring  62  depending on a desired application. Having two separate vane angle distributions for hub face  92  and shroud face  94  may help to improve vibratory′ response characteristics of turbine blades  66  by reducing High Cycle Fatigue strains at wide operating conditions. Having a single vane angle distribution for both the hub and shroud faces  92 ,  94  may be more suitable for improved stage aerodynamic performance at wide operating conditions. 
     As shown in  FIG. 11 , the thickness T V  of vanes  76  may vary along their meridional length L MV . In particular,  FIG. 11  shows a curve  120  corresponding to the thickness T V  of vanes  76  relative to the meridional length L MV  of vane  76 . As can be seen from the curve, the thickness of vanes  76  does not vary between hub and shroud faces  92 ,  94 . The thickness may reach a maximum thickness of about 5.5 min at about 20-50% of the meridional length L MV  and be thinnest at trailing edge  96 . In one embodiment, the thickness of blades  66  reaches a maximum at about 32% of the meridional length L my . Finally, maximum thickness at the leading edge may be about 0.25×T Vmax , while the maximum thickness at the trailing edge may be about 0.09×T Vn . In a similar manner to blades  66 , this smooth thickness distribution of vanes  76  may improve the manufacturability of the vanes. 
     INDUSTRIAL APPLICABILITY 
     The disclosed turbocharger may be implemented into any power system application where charged air induction is utilized. In particular, the specific geometry, blade/airfoil angle, and thickness distribution of blades  66  and vanes  76  may result in overall lower aerodynamic losses and, thus, improved performance and efficiency of turbine  32 . The uniform and well-guided flow exiting nozzle ring  62  may result in more uniform loading of nozzle ring  62  and turbine wheel  48 . This may help to reduce cyclic loading on turbine wheel  48 , extending the useful life of turbine wheel  48 . Because exhaust flow may be substantially uniform and well-guided to each blade  66 , mechanical and vibrational losses attributable to misaligned exhaust flow and turbine blade geometry may be significantly reduced. In addition, nozzle ring  62  and turbine wheel  48  may have low solidity as compared to an equivalent axial turbine stage and, thus, fewer vanes and blades. The reduction in vanes and blades may equate to a reduction in manufacturing costs. Finally, the smooth angle and thickness distribution of blades  66  and vanes  76  may allow these components to be manufactured using flank milling, which can be a cheaper alternative to other manufacturing processes. 
     It will be apparent to those skilled in the art that various modifications and variations can be made to the disclosed turbocharger. Other embodiments will be apparent to those skilled in the art from consideration of the specification and practice of the disclosed turbocharger. It is intended that the specification and examples be considered as exemplary only, with a true scope being indicated by the following claims and their equivalents.