US20150086396A1 - Turbocharger with mixed flow turbine stage - Google Patents
Turbocharger with mixed flow turbine stage Download PDFInfo
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- US20150086396A1 US20150086396A1 US14/037,836 US201314037836A US2015086396A1 US 20150086396 A1 US20150086396 A1 US 20150086396A1 US 201314037836 A US201314037836 A US 201314037836A US 2015086396 A1 US2015086396 A1 US 2015086396A1
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- Prior art keywords
- hub
- face
- shroud
- leading edge
- turbine blade
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Classifications
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D25/00—Pumping installations or systems
- F04D25/02—Units comprising pumps and their driving means
- F04D25/04—Units comprising pumps and their driving means the pump being fluid-driven
- F04D25/045—Units comprising pumps and their driving means the pump being fluid-driven the pump wheel carrying the fluid driving means, e.g. turbine blades
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D17/00—Regulating or controlling by varying flow
- F01D17/10—Final actuators
- F01D17/12—Final actuators arranged in stator parts
- F01D17/14—Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits
- F01D17/16—Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits by means of nozzle vanes
- F01D17/165—Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits by means of nozzle vanes for radial flow, i.e. the vanes turning around axes which are essentially parallel to the rotor centre line
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/141—Shape, i.e. outer, aerodynamic form
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D9/00—Stators
- F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
- F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
- F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D25/00—Pumping installations or systems
- F04D25/02—Units comprising pumps and their driving means
- F04D25/024—Units comprising pumps and their driving means the driving means being assisted by a power recovery turbine
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2220/00—Application
- F05D2220/40—Application in turbochargers
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2250/00—Geometry
- F05D2250/70—Shape
- F05D2250/75—Shape given by its similarity to a letter, e.g. T-shaped
Definitions
- the present disclosure is directed to a turbocharger and, more particularly, to a turbocharger with a mixed flow turbine stage.
- 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.
- 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.
- 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 '356 patent).
- the '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.
- the mixed flow turbine of the '356 patent may be adequate for some applications, it may still be less than optimal at wide operating conditions.
- the mixed flow turbine of the '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.
- the blade angle and thickness distributions of the mixed flow turbine shown in '356 patent are generally not smooth like a Bezier curve, which can lead to problems manufacturing the blades.
- turbocharger of the present disclosure solves one or more of the problems set forth above and/or other problems of the prior art.
- 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.
- 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.
- 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.
- 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.
- 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 ;
- FIGS. 9 , 10 , and 11 are charts associated with exemplary disclosed geometry of the nozzle ring of FIG. 3 .
- FIG. 1 illustrates a power system 10 having an engine 12 , an air induction system 14 , and an exhaust system 16 .
- engine 12 is depicted and described as a four-stroke diesel 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.
- 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 ).
- 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.
- DOC diesel oxidation catalyst
- DPF diesel particulate filter
- 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.
- turbocharger 36 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).
- 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 .
- the exhaust flow 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 ).
- 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 .
- 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 .
- 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 .
- a blade forward sweep angle ⁇ 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 ⁇ B defined by the following equation:
- ⁇ Angular coordinate, polar angle, or wrap angle
- ⁇ 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 .
- FIG. 5 illustrates a meridional view of a single blade 66 taken along the meridional length L MB .
- an R-axis defines a radial direction
- 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 ).
- leading edge 86 i.e., where exhaust enters leading edge 86 of blade 66
- an outlet flow passage or diffuser 106 adjacent to trailing edge 84
- hub curve 108 corresponding to hub face 80 and a shroud curve 110 corresponding to shroud face 82
- the relationship between hub face 80 and shroud face 82 possess unique geometric and “ruled element blade” characteristics.
- 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
- 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.
- a blade inlet cone angle ⁇ 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 .
- 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 .
- An exit deviation angle (or clip angle) ⁇ 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).
- each blade 66 may have a blade forward sweep angle ⁇ B of about 25-55°. In one embodiment, the blade forward sweep angle ⁇ B is about 47°. Blade 66 may also have a blade inlet cone angle ⁇ B of about 50-70°. In one embodiment, the blade inlet cone angle ⁇ B is about 58°.
- Blade 66 may further have a clip angle ⁇ B of about 0-14°. In one embodiment, the clip angle ⁇ 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 .
- 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.
- these geometrical features may create a blade profile that is suitable for flank milling.
- 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 .
- 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.
- the meridional blade angle ⁇ B may change along the meridional length L MB .
- FIG. 7 shows a plurality of curves corresponding to the meridional blade angle ⁇ 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 .
- the meridional blade angle ⁇ 11 at hub face 80 may be generally larger than the meridional blade angle ⁇ B at shroud face 82 (i.e., blade 66 may be more vertical at hub face 80 ).
- the meridional blade angle ⁇ B at both faces reaches a maximum at leading edge 86 and a minimum at a trailing edge 84 .
- the meridional blade angle ⁇ B generally decreases from leading edge 86 to trailing edge 84 .
- the meridional blade angle ⁇ 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 .
- the thickness T B of blades 66 may also vary along their meridional length L MB .
- 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 .
- 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 .
- the maximum thickness T Bmax may be at about 68% of the meridional length L MB . Also shown in FIG.
- the thickness of blades 66 may be substantially greater along hub face 80 than along shroud face 82 .
- a maximum thickness at the leading edge 86 may be about 0.38 ⁇ T Bmax
- 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.
- 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 ⁇ V defined by the following equation:
- ⁇ Angular coordinate, polar angle, or wrap 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 width W V may refer to a distance between hub face 92 and shroud face 94 at leading edge 86 .
- 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 ⁇ 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 .
- 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.
- 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.
- the nozzle inlet stagger angle ⁇ V of vane 76 may be about 60°-80°. In one embodiment, the nozzle inlet stagger angle ⁇ V is about 74°.
- the meridional vane angle ⁇ V of vanes 76 may change along the meridional length L MV .
- FIGS. 9 and 10 show curves corresponding to the meridional vane angle ⁇ V from leading edge 98 to trailing edge 96 for two different embodiments of nozzle ring 62 .
- the meridional vane angle ⁇ V may vary in a range of about 50-80° along its meridional length.
- the meridional vane angle ⁇ V may be substantially different along the hub and shroud faces 92 , 94 .
- a hub curve 112 may correspond to the meridional vane angle ⁇ V along the hub face 92
- a shroud curve 114 may correspond to the meridional vane angle ⁇ V along the shroud face 94 .
- both curves 112 , 114 may share a substantially S-shaped curve along the meridional length, representing a shape of the chamber of vane 76 .
- 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 .
- the meridional vane angle ⁇ V may be substantially equal along the hub and shroud faces 92 , 94 .
- FIG. 10 shows a curve 116 that corresponds to the meridional vane angle ⁇ V for both the hub and shroud faces 92 , 94 .
- 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.
- the thickness T V of vanes 76 may vary along their meridional length L MV .
- 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 .
- 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 .
- the thickness of blades 66 reaches a maximum at about 32% of the meridional length L my .
- maximum thickness at the leading edge may be about 0.25 ⁇ T Vmax
- maximum thickness at the trailing edge may be about 0.09 ⁇ T Vn .
- this smooth thickness distribution of vanes 76 may improve the manufacturability of the vanes.
- the disclosed turbocharger may be implemented into any power system application where charged air induction is utilized.
- 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.
- 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.
Abstract
Description
- The present disclosure is directed to a turbocharger and, more particularly, to a turbocharger with a mixed flow turbine stage.
- 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 '356 patent). Specifically, the '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 '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 '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 '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.
- 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.
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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 ofFIG. 2 ; -
FIG. 4 is a side-view illustration of the turbine wheel ofFIG. 3 ; -
FIG. 5 is a meridional-view illustration of an exemplary disclosed turbine blade that may be used in conjunction with the turbine wheel ofFIG. 4 ; -
FIG. 6 is a meridional-view illustration of an alternative embodiment of the turbine blade ofFIG. 5 ; -
FIGS. 7 and 8 are charts associated with exemplary disclosed geometry of the turbine blade ofFIGS. 5 ; and -
FIGS. 9 , 10, and 11 are charts associated with exemplary disclosed geometry of the nozzle ring ofFIG. 3 . -
FIG. 1 illustrates apower system 10 having anengine 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, thatengine 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 intoengine 12 for combustion. Exhaust system 16 may be configured to direct combustion exhaust fromengine 12 to the atmosphere. -
Engine 12 may include anengine block 18 that at least partially defines a plurality ofcylinders 20. A piston (not shown) may be slidably disposed within eachcylinder 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 eachcylinder 20. Eachcylinder 20, piston, and cylinder head may together at least partially define a combustion chamber. In the illustrated embodiment,engine 12 includes twelvecylinders 20 arranged in a V-configuration (i.e., a configuration having first andsecond banks engine 12 may include a greater or lesser number ofcylinders 20 and thatcylinders 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 ormore intake manifolds 30 associated withengine 12. It should be noted that air induction system 14 may includemultiple 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 ofbanks cylinders 20. Exhaust system 16 may also include at least oneturbine 32 driven by the exhaust fromexhaust manifold 34 to rotatecompressor 28 of air induction system 14.Compressor 28 andturbine 32 may together form aturbocharger 36.Turbine 32 may be configured to receive exhaust and convert potential energy in the exhaust to a mechanical rotation. After exitingturbine 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 includemultiple 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 andturbine 32 ofturbocharger 36 may be connected to each other via acommon shaft 50.Turbocharger 36 may include ahousing 40 at least partially defining compressor andturbine shrouds turbine wheels Compressor shroud 42 may include an axially-orientedinlet 52 located at a firstaxial end 54 ofturbocharger 36, and a tangentially-orientedvolute 56 located between firstaxial end 54 and a secondaxial end 58 ofturbocharger 36.Turbine shroud 44 may include avolute 60 located betweenvolute 56 and secondaxial end 58 ofturbocharger 36.Turbine shroud 44 may be configured to receive exhaust flow fromexhaust 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 anozzle ring 62.Nozzle ring 62 may be disposed downstream ofvolute 60 and be configured to accelerate exhaust gas flowing therethrough. - As
compressor wheel 46 is rotated, air may be drawn axially intoturbocharger 36 viainlet 52 and directed towardcompressor wheel 46.Blades 64 ofcompressor wheel 46 may then push the air radially outward in a spiraling fashion and into intake manifolds 30 (referring toFIG. 1 ) via an outlet volute (not shown). Similarly, as exhaust from exhaust system 16 is directed axially, radially, and tangentially inward towardturbine wheel 48, the exhaust may push againstblades 66 ofturbine wheel 48, causingturbine wheel 48 to rotate and drivecompressor wheel 46 viashaft 50. After passing throughturbine wheel 48, the exhaust flow may exit axially outward through a turbine outlet 68 located at secondaxial end 58 ofturbocharger 36 into aftertreatment system 38 (shown only inFIG. 1 ). - As illustrated in
FIG. 3 ,turbine wheel 48 may be generally disc-shaped and include a generallyannular hub 70.Blades 66 may extend outward in three dimensions fromannular hub 70.Nozzle ring 62 may be located radially upstream of turbine wheel 48 (i.e., at a periphery of turbine wheel 48). Whileturbine 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 innerannular hub 72 and an outerannular flange 74. A plurality of three-dimensional vanes 76 may be disposed between innerannular hub 72 and outerannular flange 74 to direct and accelerate exhaust flow fromvolute 60 towardblades 66 ofturbine wheel 48. - As shown in
FIG. 3 , eachblade 66 may include anairfoil 78 having a lower face (also known as a hub face) 80 that is connected tohub 70, an opposing upper face (also known as a shroud face) 82 that is oriented towards an inner surface ofshroud 44, a trailingedge 84 that is proximate to turbine outlet 68, a leadingedge 86 that is opposite to trailingedge 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 trailingedge 84 may be located closer to turbine outlet 68 than leadingedge 86. - Similarly, each
vane 76 may include a lower face (also known as a hub face) 92 that is connected tonozzle ring 62, an opposing upper face (also known as a shroud face) 94 that is oriented towards an inner surface ofshroud 44, a trailingedge 96 located proximate toturbine wheel 48, a leadingedge 98 that is opposite to trailingedge 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 trailingedge 96 may be located closer toturbine wheel 48 than leadingedge 98, -
FIG. 4 illustrates a side-view ofturbine wheel 48. For the purposes of this disclosure, a blade forward sweep angle αB ofblade 66 may refer to an angle between leadingedge 86 ofblade 66 and a base ofhub 70. A meridional length LMB ofblade 66 may refer to a meridional distance between trailing and leadingedges 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 βB defined by the following equation: -
- θ=Angular coordinate, polar angle, or wrap angle
- zm=Local meridional coordinate along the meridional length
- r=Local radial location
- βB=Local meridional blade angle
- A thickness TB may refer to a distance between low- and high-
pressure sides 88, 90 that is generally orthogonal to the camber line. A spacing SB may refer to a straight line distance between adjacent trailingedges 84 ofadjacent blades 66. A solidity ratio SRB ofblade 66 may be defined as the ratio of the meridional chord length LMB to the spacing SB (SRB=LMB/SB). -
FIG. 5 illustrates a meridional view of asingle blade 66 taken along the meridional length LMB. In the meridional plane shown inFIG. 5 , an R-axis defines a radial direction, and a Z-axis defines an axial direction along the meridional length.FIG. 5 shows aninlet flow passage 104 adjacent to leading edge 86 (i.e., where exhaust enters leadingedge 86 of blade 66), and an outlet flow passage ordiffuser 106 adjacent to trailing edge 84 (i.e., where exhaustexits trailing edge 84 of blade 66).FIG. 5 also shows ahub curve 108 corresponding tohub face 80 and ashroud curve 110 corresponding toshroud face 82, it should be noted that the relationship betweenhub 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 thathub curve 108 andshroud curve 110 are master curves and control generation of all other defining curves (e.g., intermediate curves betweenhub 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 λB may refer to an angle between the R-axis of the meridional plane and leading
edge 86 ofblade 66. An inlet hub radius r4H may refer to a distance from the Z-axis of the meridional plane to a point on thehub curve 108 at leadingedge 86. An inlet shroud radius r4S may refer to a distance from the Z-axis of the meridional plane to a point on theshroud curve 110 at leadingedge 86. An inlet width WB may refer to a distance between the point on thehub curve 108 at leadingedge 86 and the point on theshroud curve 110 at leadingedge 86. An inlet width ratio WRB may be defined as the ratio of the width WB to the meridional length LMB (WR=WB/LMB) AZ-axis offset ZB may refer to a distance between the R-axis and the point on thehub curve 108 at leadingedge 86. A non-dimensional Z-axis offset ratio ZR may be defined as the ratio of the Z-axis offset ZB to the meridional length LMB (ZRB=ZB/LMB). An exit deviation angle (or clip angle) δB may refer to an angle between trailingedge 84 ofblade 66 and the R-axis of the meridional plane. An exit hub radius r5H may refer to a distance from the Z-axis of the meridional plane to a point on thehub curve 108 at trailingedge 84. An exit shroud radius rSS may refer to a distance from the Z-axis of the meridional plane to a point on theshroud curve 110 at trailingedge 84. A turbine trim TRB may be defined by the following equation: [(r5X/r4S)2×100)]. A diffuser hub exit radius r6R TRB may refer to a distance from the Z-axis of the meridional plane to a point on thehub curve 108 at thediffuser 106. A diffuser shroud exit radius r5s may refer to a distance from the Z-axis of the meridional plane to a point on theshroud curve 110 at thediffuser 106. - The aerodynamic performance of a radial and mixed flow turbine is usually interpreted as a function of velocity ratio U/C0, where U is the blade tip speed and C0 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/C0 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/C0 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 throughturbine 32 that reduces flow misalignment (incidence) and results in improved performance and efficiency at wide operating conditions (especially at low U/C0 conditions) ofturbocharger 36. In addition, the disclosed geometry ofblade 66 increases structural integrity and manufacturability of the blades. For example, eachblade 66 may have a blade forward sweep angle αB of about 25-55°. In one embodiment, the blade forward sweep angle αB is about 47°.Blade 66 may also have a blade inlet cone angle λB of about 50-70°. In one embodiment, the blade inlet cone angle λB is about 58°.Blade 66 may further have a clip angle δB of about 0-14°. In one embodiment, the clip angle δB is about 7°. These angle ranges may help to reduce the incidence of exhaust flowing throughturbine 32 and improve vibration characteristics of theturbine 32, thereby improving aerodynamic performance and structural integrity ofturbocharger 36. - In the disclosed embodiment, the solidity ratio SRB of
blade 66 may be about 0.8-1.2, with about 10 to 17blades 66 for a giventurbine 32. In one embodiment, the solidity ratio SRB is about 1.05 for aturbine 32 housing 13 blades. The turbine trim TRB ofblade 66 may be about 50-80. In one embodiment, the turbine trim TRB is about 59. The width ratio WRB ofblade 66 may be about 0.2-0.42. In one embodiment, the width ratio WRB is about 0.29. The Z-axis offset ratio ZRB, ofblade 66 may be about 0.07-0.20. In one embodiment, the Z-axis offset ratio ZRB is about 0.13. Each of these geometrical features may help to improve aerodynamic performance and structural integrity ofblades 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 aninlet flow passage 104 adjacent to leading edge 86 (i.e., where exhaust enters leadingedge 86 of blade 66), and an outlet flow passage ordiffuser 106 adjacent to trailing edge 84 (i.e., where exhaustexits trailing edge 84 of blade 66). The disclosed geometry ofinlet flow passage 104 has been selected to provide a desired aerodynamic flow guidance and uniformity into the turbineblade 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) ofturbocharger 36. Theoutlet flow passage 106 which acts like adiffuser 106 may have a diffuser ratio at hub (r5h/r6h) from 1.15 to 1.55. In one embodiment, diffuser ratio at hub is 1.35. Theoutlet flow passage 106 which acts like adiffuser 106 may have a diffuser ratio at shroud (r6s/r5s) from 1.02 to 1.10. In one embodiment, diffuser ratio at shroud is 1.07. -
FIG. 6 shows an alternative embodiment ofblade 66. In this embodiment, leadingedge 86 ofblade 66 is substantially concave rather than being substantially straight as shown in the embodiment ofFIG. 5 . It is contemplated that having a concaveleading 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 βB may change along the meridional length LMB. Specifically,FIG. 7 shows a plurality of curves corresponding to the meridional blade angle β3 betweenhub curve 108 andshroud curve 110. It should be noted that each curve betweenhub curve 108 andshroud curve 110 may correspond with an intermediate layer ofblade 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 β11 athub face 80 may be generally larger than the meridional blade angle βB at shroud face 82 (i.e.,blade 66 may be more vertical at hub face 80). In addition, the meridional blade angle βB at both faces reaches a maximum at leadingedge 86 and a minimum at a trailingedge 84. In other words, the meridional blade angle βB generally decreases from leadingedge 86 to trailingedge 84. As also shown inFIG. 7 , the meridional blade angle βB may vary between about −5° and 30° at leadingedge 86, and vary between about −40° and −80° at trailingedge 84. This blade angle distribution may help to reduce aerodynamic losses and, thus, improve performance and efficiency ofturbocharger 36. - As shown in
FIG. 8 , the thickness TB ofblades 66 may also vary along their meridional length LMB. In particular,FIG. 8 shows a plurality of curves corresponding to the thickness TB ofblades 66 between the hub and shroud faces 80, 82 relative to the meridional length LMB ofblades 66. As can be seen from the plurality of curves, the thickness ofblades 66 may reach a maximum thickness TBmax of about 10 mm at about 60-80% of the meridional length 140 and be thinnest at trailing and leadingedges FIG. 8 , the thickness ofblades 66 may be substantially greater alonghub face 80 than alongshroud face 82. Finally, a maximum thickness at theleading edge 86 may be about 0.38×TBmax, while a maximum thickness at the trailingedge 84 may be about 0.61×TBmax. This smooth Bezier curve thickness distribution ofblades 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 ofvanes 76 ofnozzle ring 62 will now be discussed. For the purposes of this disclosure, a meridional length LMV may refer to a meridional distance between trailing and leadingedges vanes 76 along a camber line passing through a lengthwise center of the vanes. Similar toblades 66,vanes 76 may also curve along their lengths, each forming a corresponding meridional vane angle βV defined by the following equation: -
- θ=Angular coordinate, polar angle, or wrap angle
- zm=Local meridional coordinate along the meridional length
- r=Local radial location
- βV=Local meridional vane angle
- A thickness TV may refer to a distance between high- and low-
pressure sides vane 76. A chord length LCV may refer to a straight line distance between trailing and leadingedges vanes 76. A spacing SV may refer to a straight line distance between adjacent trailingedges 96 ofadjacent vanes 76. A solidity ratio SRV may be defined as the ratio of the chord length LCV to the spacing SV (SRV=LCV/SV). A width WV may refer to a distance betweenhub face 92 and shroud face 94 at leadingedge 86. A width ratio WRV may be defined as the ratio of the width WV to the chord length LCV (WRV=WV/LCV). A blade inlet shroud tip radius r1 may refer to a distance from a center ofturbine wheel 48 to leadingedge 86 ofblade 66 atshroud face 82. A vane leading edge radius r2 may refer to a distance from the center ofturbine wheel 48 to leadingedge 98 ofvane 76 atshroud face 82. A vane inlet radius ratio My may be defined as the ratio of the vane leading edge radius r2 to the blade inlet shroud tip radius r1. A nozzle inlet stagger angle φV may refer to an angle between the chord length Lcv and the vane leading edge radius r2. A vane trailing edge radius r3 may refer to a distance from the center ofturbine wheel 48 to trailingedge 96 ofvane 76 atshroud face 82. A vane exit radius ratio ERV may be defined as the ratio of the vane trailing edge radius r3 to the blade inlet shroud tip radius r1. - Similar to
blades 66, the disclosed geometry ofvanes 76 has been selected to provide desired aerodynamic flow angles with improved flow uniformity at an exit ofnozzle ring 62, increased structural integrity of the vanes, and low torque loading of thevanes 76. For example, eachvane 76 may have a solidity ratio SRV of about 0.7-1.2, with about 13 to 25vanes 76 included aroundnozzle ring 62. In one embodiment, the solidity ratio is about 1.11, with 23 blades included aroundnozzle ring 62. The width ratio WRV ofvane 76 may be about 0.2-0.40. In one embodiment, the width ratio WRV is about 0.23. The vane inlet radius ratio IR ofvane 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 ofvane 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 φV ofvane 76 may be about 60°-80°. In one embodiment, the nozzle inlet stagger angle φV is about 74°. Each of these geometrical features may help to reduce aerodynamic losses, reduce vane torque loading, and improve the structural integrity ofvanes 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 βV ofvanes 76 may change along the meridional length LMV. Specifically,FIGS. 9 and 10 show curves corresponding to the meridional vane angle βV from leadingedge 98 to trailingedge 96 for two different embodiments ofnozzle ring 62. In each of these embodiments, the meridional vane angle βV 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 βV 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 βV along thehub face 92, and ashroud curve 114 may correspond to the meridional vane angle βV along theshroud face 94. In this embodiment, bothcurves 112, 114 may share a substantially S-shaped curve along the meridional length, representing a shape of the chamber ofvane 76. However, the hub curve 112 may be substantially greater than theshroud curve 114 at each point along the meridional length between leadingedge 98 and trailingedge 96. - In a second embodiment shown in
FIG. 10 , the meridional vane angle γV may be substantially equal along the hub and shroud faces 92, 94. Thus, only one curve is shown.FIG. 10 shows acurve 116 that corresponds to the meridional vane angle βV 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 leadingedge 98 to trailingedge 96, shown here asinclination curve 118. Both of the above embodiments ofvanes 76 may be used withnozzle ring 62 depending on a desired application. Having two separate vane angle distributions forhub face 92 and shroud face 94 may help to improve vibratory′ response characteristics ofturbine 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 TV ofvanes 76 may vary along their meridional length LMV. In particular,FIG. 11 shows a curve 120 corresponding to the thickness TV ofvanes 76 relative to the meridional length LMV ofvane 76. As can be seen from the curve, the thickness ofvanes 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 LMV and be thinnest at trailingedge 96. In one embodiment, the thickness ofblades 66 reaches a maximum at about 32% of the meridional length Lmy. Finally, maximum thickness at the leading edge may be about 0.25×TVmax, while the maximum thickness at the trailing edge may be about 0.09×TVn. In a similar manner toblades 66, this smooth thickness distribution ofvanes 76 may improve the manufacturability of the vanes. - 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 andvanes 76 may result in overall lower aerodynamic losses and, thus, improved performance and efficiency ofturbine 32. The uniform and well-guided flow exitingnozzle ring 62 may result in more uniform loading ofnozzle ring 62 andturbine wheel 48. This may help to reduce cyclic loading onturbine wheel 48, extending the useful life ofturbine wheel 48. Because exhaust flow may be substantially uniform and well-guided to eachblade 66, mechanical and vibrational losses attributable to misaligned exhaust flow and turbine blade geometry may be significantly reduced. In addition,nozzle ring 62 andturbine 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 ofblades 66 andvanes 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.
Claims (32)
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US14/037,836 US20150086396A1 (en) | 2013-09-26 | 2013-09-26 | Turbocharger with mixed flow turbine stage |
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