TURBOCHARGER WITH MIXED FLOW TURBINE STAGE
A turbocharger is disclosed for use with an engine. 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 leading edge may be substantially straight or substantially concave in a meridional plane. An angle between a base of the hub and the leading edge may be about 25-55 degrees. 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.
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The present disclosure is directed to a turbocharger and, more particularly, to a turbocharger with a mixed flow turbine stage.
BACKGROUNDInternal 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.
SUMMARYIn 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.
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
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
As illustrated in
As shown in
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,
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 trailing edges 84 of adjacent blades 66. A solidity ratio SRB of blade 66 may be defined as the ratio of the meridional chord length LMB to the spacing SB (SRB=LMB/SB).
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 of blade 66. An inlet hub radius r4H 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 r4S 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 WB 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 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 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 ZB to the meridional length LMB (ZRB=ZB/LMB). An exit deviation angle (or clip angle) δB may refer to an angle between trailing edge 84 of blade 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 the hub curve 108 at trailing edge 84. An exit shroud radius rSS 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 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 the hub curve 108 at the diffuser 106. A diffuser shroud exit radius r5s 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/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 through turbine 32 that reduces flow misalignment (incidence) and results in improved performance and efficiency at wide operating conditions (especially at low U/C0 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 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 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 SRB 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 SRB is about 1.05 for a turbine 32 housing 13 blades. The turbine trim TRB of blade 66 may be about 50-80. In one embodiment, the turbine trim TRB is about 59. The width ratio WRB of blade 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, of blade 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 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,
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,
As shown in
Referring back to
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 100, 102 that is generally orthogonal to the camber line of vane 76. A chord length LCV may refer to a straight line distance between trailing and leading edges 96, 98 of vanes 76. A spacing SV may refer to a straight line distance between adjacent trailing edges 96 of adjacent 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 between hub face 92 and shroud face 94 at leading edge 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 of turbine wheel 48 to leading edge 86 of blade 66 at shroud face 82. A vane leading edge radius r2 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 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 of turbine wheel 48 to trailing edge 96 of vane 76 at shroud 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 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 SRV 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 WRV of vane 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 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 φV of vane 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 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 βV of vanes 76 may change along the meridional length LMV. Specifically,
In a first embodiment shown in
In a second embodiment shown in
As shown in
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.
Claims
1. A turbine blade for a turbocharger, comprising:
- 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,
- wherein an angle between the turbine wheel hub and the leading edge is about 25-55 degrees.
2. The turbine blade of claim 1, wherein a solidity ratio of the turbine blade ranges from about 0.8 to 1.2 between the hub face and the shroud face.
3. The turbine blade of claim 1, wherein an angle between a radial axis in a meridional plane of the turbine blade and the leading edge in the meridional plane is about 50-70 degrees.
4. The turbine blade of claim 1, wherein an angle between a radial axis in a meridional plane of the turbine blade and the trailing edge in the meridional plane is about 0-14 degrees.
5. The turbine blade of claim 1, wherein a ratio between an inlet width of the leading edge in a meridional plane of the turbine blade and a meriodional distance between the leading edge and the trailing edge ranges from about 0.20 to 0.42 between the hub face and the shroud face.
6. The turbine blade of claim 1, wherein a ratio between a Z-axis offset in a meridional plane of the turbine blade and a meriodional distance between the leading edge and the trailing edge ranges from about 0.07 to 0.20 between the hub face and the shroud face.
7. The turbine blade of claim 1, wherein the turbine blade has a blade angle that changes along its meridional length from about −80° to 30°.
8. The turbine blade of claim 7, wherein the blade angle decreases from the hub face to the shroud face.
9. The turbine blade of claim 7, wherein the blade angle is greatest for the turbine blade at the leading edge.
10. The turbine blade of claim 9, wherein the blade angle ranges from about −5 to 30 degrees between the hub face and the shroud face at the leading edge.
11. The turbine blade of claim 7, wherein the blade angle is lowest for the turbine blade at the trailing edge.
12. The turbine blade of claim 11, wherein the blade angle ranges from about −40 to −80 degrees between the hub face and the shroud face at the trailing edge.
13. The turbine blade of claim 1, wherein a thickness of the turbine blade varies along a meridional length of the turbine blade and decreases from the hub face to the shroud face.
14. The turbine blade of claim 13, wherein the thickness is greatest at about 60-80% of the meridional length of the turbine blade.
15. The turbine blade of claim 14, wherein a maximum thickness at the leading edge is about 0.38 times the greatest thickness along the meridional length.
16. The turbine blade of claim 14, wherein a maximum thickness at the trailing edge is about 0.61 times the greatest thickness along the meridional length.
17. The turbine blade of claim 1, wherein the leading edge is substantially straight.
18. The turbine blade of claim 1, wherein the leading edge is substantially concave.
19. A nozzle ring for a turbocharger, comprising:
- a ring-shaped generally flat plate having: an inner annular hub; and an outer annular flange radially spaced apart from the inner annular hub; and
- a plurality of vanes disposed between the inner annular hub and the outer annular flange,
- wherein a camber of each of the plurality of vanes is generally S-shaped along its meridional length from a leading edge to a trailing edge of each of the plurality of vanes.
20. The nozzle ring of claim 19, wherein the S-shaped camber generally inclines along its meridional length from the leading edge to the trailing edge.
21. The nozzle ring of claim 19, wherein a vane angle of each of the plurality of vanes changes along its meridional length in a range from about 50 to 80 degrees.
22. The nozzle ring of claim 19, wherein:
- each of the plurality of vanes includes a hub face and a shroud face; and
- a vane angle along the hub face is substantially different from a vane angle along the shroud face.
23. The nozzle ring of claim 19, wherein:
- each of the plurality of vanes includes a hub face and a shroud face; and
- a vane angle along the hub face is substantially equal to a vane angle along the shroud face.
24. The nozzle ring of claim 19, wherein a thickness of each of the plurality of vanes varies along its meridional length, the thickness being greatest at about 20-50% of the meridional length.
25. The nozzle ring of claim 24, wherein a maximum thickness at the leading edge is about 0.25 times the greatest thickness along the meridional length.
26. The nozzle ring of claim 24, wherein a maximum thickness at the trailing edge is about 0.09 times the greatest thickness along the meridional length.
27. The nozzle ring of claim 19, wherein:
- each of the plurality of vanes includes a hub face and a shroud face; and
- a solidity ratio of each of the plurality of vanes ranges from about 0.7 to 1.2 between the hub face and the shroud face.
28. The nozzle ring of claim 27, wherein a ratio between a width of each of the plurality of vanes and a chord distance between the leading edge and the trailing edge ranges from about 0.20 to 0.40 between the hub face and the shroud face.
29. The nozzle ring of claim 28, wherein a ratio between a distance from a center of a hub associated with the turbocharger to the leading edge of each of the plurality of vanes at the shroud face, and a distance from the center of the hub to a leading edge of a blade associated with the turbocharger at the shroud face ranges from about 1.3 to 1.5 between the hub face and the shroud face.
30. The nozzle ring of claim 29, wherein a ratio between a distance from the center of the hub to the trailing edge of each of the plurality of vanes at the shroud face, and the distance from the center of the hub to the leading edge of the blade at the shroud face ranges from about 1.05 to 1.3 between the hub face and the shroud face.
31. The nozzle ring of claim 30, wherein an angle between a line drawn between the center of the hub to the leading edge of each of the plurality of vanes at the shroud face, and a line drawn between the leading edge and the trailing edge is about 60-80 degrees.
32. A turbocharger, comprising:
- a housing at least partially defining a compressor shroud and a turbine shroud;
- a compressor wheel disposed within the compressor shroud;
- a shaft connected to the compressor wheel;
- a turbine wheel disposed within the turbine shroud and connected to an end of the shall opposite the compressor wheel, the turbine wheel including: a generally annular hub; and a plurality of blades disposed radially around the annular hub, each of the plurality of blades including an airfoil having a hub face connected to the annular 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, wherein an angle between a base of the annular hub and the leading edge is about 25-55 degrees; and
- a nozzle ring including: 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, wherein a camber of each of the plurality of vanes is generally S-shaped along its meridional length from a leading edge to a trailing edge of each of the plurality of vanes.
Type: Application
Filed: Sep 26, 2013
Publication Date: Mar 26, 2015
Applicant: Electro-Motive Diesel Inc. (LaGrange, IL)
Inventor: Shakeel Nasir (Willowbrook, IL)
Application Number: 14/037,836
International Classification: F04D 25/04 (20060101); F01D 9/04 (20060101); F01D 5/14 (20060101);