CN104514631A - Turbocharger with mixed flow turbine stage - Google Patents
Turbocharger with mixed flow turbine stage Download PDFInfo
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- CN104514631A CN104514631A CN201410494389.XA CN201410494389A CN104514631A CN 104514631 A CN104514631 A CN 104514631A CN 201410494389 A CN201410494389 A CN 201410494389A CN 104514631 A CN104514631 A CN 104514631A
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- Prior art keywords
- hub
- blade
- face
- fin
- guard shield
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Classifications
-
- 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
- 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
- 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
Abstract
A turbocharger for use with an engine is disclosed. 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 radially disposed about 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 toward the turbine shroud, a trailing edge, and a leading edge opposite the trailing edge. The angle between the base and the leading edge of the hub may be about 25-55 degrees. The turbocharger may also include a nozzle ring having an annular, generally flat disc located at the periphery of the turbine wheel and a plurality of vanes radially disposed about an upper surface of the disc. The arc of each of the plurality of vanes may be generally S-shaped along a meridional length thereof from a leading edge to a trailing edge of each of the plurality of vanes.
Description
Technical field
The present invention relates to a kind of turbosupercharger, more specifically, relate to the turbosupercharger with mixed flow turbine level.
Background technique
Explosive motor, such as, motor as diesel engine, petrol engine and gaseous fuel-driven is supplied to the mixture of air and fuel with subsequent combustion in the motor producing machine power output.In order to increase the Power output that combustion process thus produces, motor can be equipped with turbosupercharging gas handling system.
Turbosupercharging gas handling system comprises turbosupercharger, and it uses from the exhaust of motor to compress the air of inflow engine, thus forces more air to enter engine chamber compared with otherwise sucking the motor of firing chamber.The air supply of this increase allows fuel of increasing supply, thus increases Power output.Turbosupercharged engine is usually than not having turbo charged same motor to produce more power.
Conventional turbosupercharger comprises turbine cylinder and is arranged on center in housing and the turbine wheel of the compressor impeller be connected with rotation by exhaust gas drive.Exhaust is pushed against and is connected on the blade of turbine wheel, rotates to make turbine wheel.In some applications, being arranged on the fin be connected on the nozzle ring of turbine wheel makes exhaust accelerate through blade.The blade of fin and/or turbo machine can axially, directing exhaust gas in radial and tangential direction.
Mixed flow turbine is regarded as the cross-over design between radial and axial turbo machine usually.Authorize on March 6th, 2012 in U.S.'s No. 8128356 patent (' 356 patent) of Higashimori and disclose a kind of exemplary mixed flow turbine.Particularly, ' 356 patent describes a kind of mixed flow turbine with blade and scroll, the leading edge profile being positioned at upstream side of blade is shaped as towards the convex of upstream side, and scroll is the space of the blade upstream formed by the housing with the radially outer edge guard shield covering blade.Process fluid is supplied to hub and guard shield, the intake duct of shroud and hub side intake duct place substantially vertically, radial and tangential direction flowing.The shape of the leading edge of blade is designed to reduce impact loss.
Although perhaps the mixed flow turbine of ' 356 patent is enough to some application, still be apparent not enough compared with optimum under wide in range operational condition.Especially, the mixed flow turbine of ' 356 patent can guide the mixed flow of uneven and bad guiding to pass through turbine stage under wide in range operational condition, due to flowing and turbine bucket misalignment (the high angle of attack), this can cause the aerodynamic efficiency of high energy losses, reduction in operation, and the machinery increased on turbo machine or vibration stress (or strain).In addition, the angle of the mixed flow turbine blade shown in the patent of ' 356 and thickness distribution are usually level and smooth unlike Bezier curve, and this can cause the problem of vane manufacturing.
Turbosupercharger of the present invention solves one or more problem presented above and/or other problem of the prior art.
Summary of the invention
On the one hand, the present invention relates to a kind of turbosupercharger.This turbosupercharger can comprise the housing limiting compressor protective cover and turbomachine shroud at least in part.This turbosupercharger can also comprise the compressor impeller be arranged in compressor protective cover, be connected to the axle of compressor impeller and be arranged in turbomachine shroud and be connected to the turbine wheel of one end relative with compressor impeller of axle.Turbine wheel can have the hub and multiple blade around hub diameter to setting that are roughly annular.Described in each, multiple blade can comprise aerofoil, this aerofoil have be connected to hub hub face, relative with hub face and be oriented towards the guard shield face of turbomachine shroud, trailing edge and the leading edge relative with trailing edge.Angle between the pedestal of hub and leading edge can be about 25-55 degree.Leading edge can be straight substantially or be recessed substantially in meridional plane.Turbosupercharger can also comprise nozzle ring, nozzle ring have be positioned at annular around turbine wheel be roughly flat dish and fin that multiple radial direction of the upper surface around dish is arranged.The arc of multiple fin described in each can be substantially S shape along its meridian length from leading edge to the trailing edge of fin multiple described in each.
Second aspect, the present invention relates to a kind of turbine bucket of turbosupercharger.Turbine bucket can comprise aerofoil, aerofoil have the turbine wheel hub being connected to turbosupercharger hub face, relative with hub face and be oriented towards the guard shield face of the turbomachine shroud of turbosupercharger, trailing edge and the leading edge relative with trailing edge.Angle between the pedestal of turbine wheel and leading edge can be about 25-55 degree.Leading edge can be straight substantially or be recessed substantially in meridional plane.
The third aspect, the present invention relates to a kind of nozzle ring for turbosupercharger.What this nozzle ring can comprise the ring-type with inner annular hub is roughly flat dish, and with inner annular hub diameter to the exterior annular flange separated.Nozzle ring can also comprise the multiple fins be arranged between inner annular hub and exterior annular flange.The arc of multiple fin described in each can be substantially S shape along its meridian length from leading edge to the trailing edge of each multiple fin described.
Accompanying drawing explanation
Fig. 1 is the schematic diagram of the power system of exemplary disclosure;
Fig. 2 is the schematic cross-section of the turbosupercharger of the exemplary disclosure that can be combined with the power system in Fig. 1;
Fig. 3 is the turbine wheel of exemplary disclosure and the schematic diagram directly perceived of nozzle ring that can be combined with the turbosupercharger in Fig. 2;
Fig. 4 is the side schematic view of turbine wheel in Fig. 3;
Fig. 5 is the meridian plane view of the turbine bucket of the exemplary disclosure that can be combined with the turbine wheel in Fig. 4;
Fig. 6 is the meridian view of the alternate embodiments of turbine bucket in Fig. 5;
Fig. 7 and Fig. 8 is the chart be associated with the geometrical construction of the exemplary disclosure of turbine bucket in Fig. 5; And
Fig. 9, Figure 10 and Figure 11 are the chart be associated with the geometrical construction of the exemplary disclosure of nozzle ring in Fig. 3.
Embodiment
Fig. 1 illustrates a kind of power system 10, and it has motor 12, gas handling system 14 and vent systems 16.For the purposes of the present invention, motor 12 is described and is described as four-cycle diesel engine.But, it will be appreciated by those of skill in the art that motor 12 can be the combustion engine of other type any, such as, the such as motor of two-stroke or four-stroke gasoline or gaseous fuel-driven.Gas handling system 14 can be set to guide the mixture of air or air and fuel to enter motor 12 to burn.Vent systems 16 can be set to guide burning and gas-exhausting from motor 12 to air.
Motor 12 can comprise the cluster engine 18 limiting multiple cylinder 20 at least in part.Piston (not shown) can be slidably disposed in each cylinder 20, to-and-fro motion between top dead center position and bottom dead center position, and cylinder head (not shown) can be associated with each cylinder 20.Each cylinder 20, piston and cylinder head can limit firing chamber together at least in part.In the exemplary embodiment, motor 12 comprises 12 cylinders 20 arranged with V-type structure (that is, having the first and second rows 22,24 of cylinder or the structure of row).But it is contemplated that as required, motor 12 can comprise the cylinder 20 of more or less quantity, and cylinder 20 can be arranged to structure in upright arrangement, opposed pistons structure or other structure.
Inter alia, gas handling system 14 can comprise at least one compressor 28, and it can turn to the compressor of fixed geometirc structure, variable-geometry compressor or can admission of air by the compressor of air compressing to other type any of the stress level expected by tool.The bootable air of compressor 28 is to the one or more intake manifold 30 associated with motor 12.It should be noted that gas handling system 14 can comprise multiple compressor 28, it is arranged as the composite construction of series configuration, parallel construction or series configuration/parallel construction.
Inter alia, vent systems 16 can comprise the gas exhaust manifold 34 of one or two row 22,24 being connected to cylinder 20.Vent systems 16 also can comprise at least one turbo machine 32 by the exhaust gas drive from gas exhaust manifold 34, with the compressor 28 of rotating inlet system 14.Compressor 28 and turbo machine 32 can form turbosupercharger 36 jointly.Turbo machine 32 can be set to receive exhaust and the potential energy converting and energy in exhaust be rotated for mechanical.After leaving turbo machine 32, exhaust is discharged in air by after-treatment system 38, after-treatment system 38 can comprise such as hydrocarbon doser, diesel oxidation catalyst (DOC), diesel particulate filter (DPF) and/or other any treatment device known in the art, if necessary.It should be noted that vent systems 16 can comprise multiple turbo machine 32, it can by expecting the composite construction being arranged to series configuration, parallel construction or series/parallel.
As shown in Figure 2, the compressor 28 of turbosupercharger 36 and turbo machine 32 are connected to each other by common shaft 50.Turbosupercharger 36 can comprise housing 40, and it limits the compressor protective cover 42 and turbomachine shroud 44 that can hold corresponding compressor impeller 46 and turbine wheel 48 at least in part.Compressor protective cover 42 can comprise the volute (volute) 56 of the entrance 52 of the axial orientation of the first axial end 54 in turbosupercharger 36 and the tangential orientation between first axial end 54 and the second axial end 58 of turbosupercharger 36.Turbomachine shroud 44 can comprise the volute 60 between the volute 56 and the second axial end 58 of turbosupercharger 36.Turbomachine shroud 44 can be set at the exhaust stream of solenoid inlet (not shown) reception from the tangent direction of gas exhaust manifold 34.Volute 60 can at three direction directing exhaust gas streams: axially (along spin axis X), radially-inwardly (along volute radius) and tangential (around spin axis X) towards and pass through nozzle ring 62.Nozzle ring 62 can be arranged on the downstream of volute 60 and can accelerate exhaust from wherein flowing through.
Because compressor impeller 46 is rotated, air can be inhaled in turbosupercharger 36 via entrance 52 vertically, and is guided by towards compressor impeller 46.Then the blade 64 of compressor impeller 46 by air radially outward release in a spiral manner, and can enter intake manifold 30 (see Fig. 1) via outlet volute (not shown).Similarly, because the exhaust from vent systems 16 is guided by axial, radial and tangential inwardly turbine wheel 48, exhaust can push against the blade 66 of turbine wheel 48, causes turbine wheel 48 to rotate and drives compressor impeller 46 by axle 50.After passing turbine wheel 48, exhaust stream can axially outwards leave, and enters after-treatment system 38 (being only shown in Fig. 1) by the turbine outlet 68 of the second axial end 58 being positioned at turbosupercharger 36.
As shown in Figure 3, turbine wheel 48 can be cardinal principle plate-like and comprises the hub 70 being roughly annular.Blade 66 can stretch out from annular hub 70 with three dimensions.Nozzle ring 62 can be positioned in the footpath of turbine wheel 48 upstream (that is, in the periphery of turbine wheel 48).When turbine wheel 48 rotates with sense of rotation R, nozzle ring 62 can be static.Nozzle ring 62 can be cast substantially, and comprises inner annular hub 72 and exterior annular flange 74.Multiple three-dimensional fin 76 can be arranged between inner annular hub 72 and exterior annular flange 74, flows to guide and to accelerate the blade 66 be vented from volute 60 towards turbine wheel 48.
As shown in Figure 3, each blade 66 can comprise aerofoil 78, its have be connected to (be also referred to as hub face) below hub 70 80, be oriented towards the internal surface of guard shield 44 relative above (being also referred to as guard shield face) 82, the trailing edge 84 close to turbine outlet 68, the leading edge 86 relative with trailing edge 84, high pressure side (being also referred to as on the pressure side) 88 and relative low voltage side (being also referred to as suction side) 90.It is envisioned that trailing edge 84 can than leading edge 86 closer to turbine outlet 68.
Similarly, each fin 76 can comprise be connected to (be also referred to as hub face) below nozzle ring 62 92, be oriented towards the internal surface of guard shield 44 relative above (being also referred to as guard shield face) 94, the trailing edge 96 close to turbine wheel 48, the leading edge 98 relative with trailing edge 96, high pressure side (being also referred to as on the pressure side) 100 and relative low voltage side (being also referred to as suction side) 102.It is envisioned that trailing edge 96 can than leading edge 98 closer to turbine wheel 48.
Fig. 4 shows the side view of turbine wheel 48, for the purposes of the present invention, and the blade sweepforward angle α of blade 66
bthe angle between the leading edge 86 of blade 66 and the pedestal of hub 70 can be referred to.The meridian length L of blade 66
mBcan refer between the trailing edge 84 of blade 66 and leading edge 86 along through the curved line of the longitudinal center of blade meridional distance from.Blade 66 can along its curved in length, each meridian blade angles forming correspondence
b, it is defined by equation below:
θ=angular coordinates, polar angle or cornerite
Z
m=along the local meridian coordinate of meridian length
R=local radial position
β
b=local meridian blade angle
Thickness T
bthe usual distance perpendicular to curved line between low voltage side 88 and high pressure side 90 can be referred to.Interval S
bcrow flight distance between the adjacent trailing edge 84 that can refer to adjacent blades 66.The solidity ratio SR of blade 66
bmeridian chord length L can be defined as
mBand interval S
bratio (SR
b=L
mB/ S
b).
Fig. 5 illustrates that individual blade 66 is along meridian length L
mBmeridian view.In meridional plane in Figure 5, R axis limits radial direction, and z axis defines the axial direction along meridian length.Fig. 5 illustrates the entrance channel 104 (that is, exhaust enters the position of the leading edge 86 of blade 66) adjacent to leading edge 86 and the outlet flow adjacent to trailing edge 84 or diffuser 106 (that is, the position of the trailing edge 84 of blade 66 is left in exhaust).Fig. 5 also illustrates the hub curve 108 corresponding to hub face 80 and the guard shield curve 110 corresponding to guard shield face 82.It should be noted that the relation between hub face 80 and guard shield face 82 has unique geometric properties and " straight burr element blade (ruled element blade) " feature.In straight burr element blade, Angle Position is limited by the straight line drawn between the point along the span location between hub face 80 and guard shield face 82 in three dimensions.Should also be noted that hub curve 108 and guard shield curve 110 are principal curve and control the generation of all curves (that is, the intermediate curve between hub curve 108 and guard shield curve 110) that other limits.The amendment of hub curve 108 and/or guard shield curve 110 can cause the subsequent modification of intermediate curve.
For the purposes of the present invention, blade inlet cone angle λ
bthe angle between the R axis of the meridional plane of blade 66 and leading edge 86 can be referred to.Entrance hub radius r
4Hthe distance of the point being positioned at leading edge 86 from the z axis of meridional plane to guard shield curve 108 can be referred to.Entrance shield radius r
4Sthe distance of the point being positioned at leading edge 86 from the z axis of meridional plane to guard shield curve 110 can be referred to.Throat width W
bthe distance on the point of leading edge 86 and guard shield curve 110 between the point of leading edge 86 on hub curve 108 can be referred to.Throat width compares WR
bwidth W can be defined as
bwith meridian length L
mBratio (WR=W
b/ L
mB).Z axis skew Z
bthe distance between the point of leading edge 86 on R axis and hub curve 108 can be referred to.Nondimensional z axis deviation ratio ZR can be defined as z axis skew Z
bwith meridian length L
mBratio (ZR
b=Z
b/ L
mB).Outlet deviation angle (or angle) δ
bthe angle between the trailing edge 84 of blade 66 and the R axis of meridional plane can be referred to.Outlet hub radius r
5Hthe distance of the some z axis of meridional plane to hub curve 108 being positioned at trailing edge 84 can be referred to.Outlet shield radius r
5Sthe distance of the point being positioned at trailing edge 84 from the z axis of meridional plane to guard shield curve 110 can be referred to.TR pruned by turbo machine
bcan be defined by following formula: [(r
5s/ r
4s)
2× 100)].Diffuser hub exit radius r
6Hthe distance of the point being positioned at diffuser 106 from the z axis of meridional plane to hub curve 108 can be referred to.Diffuser guard shield exit radius r
5Sthe distance of the point being positioned at diffuser 106 from the z axis of meridional plane to guard shield curve 110 can be referred to.
Aerodynamic quality that is radial and mixed flow turbine is understood to velocity ratio U/C usually
ofunction, wherein U is blade tip velocity, C
obe constant entropy speed, got by the desired expansion of the gas by the pressure ratio equal with turbo machine.Because turbosupercharger often needs at low U/C
ooperational condition (or high expansion ratio condition of constant tip speed) under operation, so need the design of a kind of effective turbine stage with at these low U/C
othere is when operating under condition low aerodynamic losses (such as, impact loss).The geometrical construction of blade 66 of the present invention is selected to provide the Aerodynamic Flows uniformity by the expectation of turbo machine 32 and guidance quality, reduces flowing deviation (impact) and makes turbosupercharger 36 at wide in range operational condition (especially low U/C
ocondition) under there is performance and the efficiency of improvement.In addition, the geometrical construction of blade 66 of the present invention increases structural integrity and the manufacturability of blade.Such as, each blade 66 can have the blade sweepforward angle α of about 25-55 °
b.In one embodiment, this blade sweepforward angle α
bit can be about 47 °.Blade 66 also can have the blade inlet cone angle λ of about 50-70 °
b.In one embodiment, this blade inlet cone angle λ
bit can be about 58 °.Blade 66 can also have the angle δ of about 0-14 °
b.In one embodiment, this angle δ
bfor about 7 °.These angular regions can contribute to reducing the impact of exhaust stream through turbo machine 32, and improve the vibration characteristics of turbo machine 32, thus improve aerodynamic quality and the structural integrity of turbosupercharger 36.
In embodiments of the present invention, for given turbo machine 32, there are about 10 to 17 blades 66, the solidity ratio SR of blade 66
bcan be about 0.8-1.2.In one embodiment, for the turbo machine 32, solidity ratio SR of accommodation 13 blades
bfor about 1.05.TR pruned by the turbo machine of blade 66
bcan be about 50-80.In one embodiment, TR pruned by this turbo machine
bfor about 59.The width of blade 66 compares WR
bcan be about 0.2-0.42.In one embodiment, this width compares WR
bfor about 0.29.The z axis deviation ratio ZR of blade 66
bcan be about 0.07-0.20.In one embodiment, z axis deviation ratio ZR
bfor about 0.13.Each these geometric properties all can contribute to aerodynamic quality and the structural integrity of improving blade 66, simultaneously with respect to being conducive to the smoothed curve improving manufacturability.Especially, these geometric properties can form the blade profile being suitable for Flank machining.
As mentioned above, (namely Fig. 5 shows the entrance channel 104 adjacent with leading edge 86, exhaust enters the position of the leading edge 86 of blade 66) and the outlet flow adjacent with trailing edge 84 or diffuser 106 (that is, being vented the position of the trailing edge 84 leaving blade 66).The geometrical construction of entrance channel 104 of the present invention is selected as providing the Aerodynamic Flows guidance quality of expectation and uniformity to enter turbine blade leading edge 86, and this also reduces flowing deviation (impact) and makes turbosupercharger 36 at wide in range operational condition (especially low U/C
ocondition) under there is performance and the efficiency of improvement.The effect of outlet flow 106 is similar to diffuser 106, and it can have the diffusivity (r from 1.15 to 1.55 at hub place
5h/ r
6h).In one embodiment, diffusivity is 1.35 at hub place.The outlet flow 106 that effect is similar to diffuser 106 can have the diffusivity (r from 1.02 to 1.10 at guard shield place
6s/ r
5s).In one embodiment, be 1.07 in guard shield place diffusivity.
Fig. 6 shows a kind of alternate embodiments of blade 66.In this embodiment, the leading edge 86 of blade 66 be roughly spill instead of being roughly as shown in mode of execution in Fig. 5 straight.Can expect there is the flowing that spill leading edge 86 can contribute to improving further in some applications under wide in range operational condition and aim at.
In order to improve manufacturability and the aerodynamic quality of blade 66 further, meridian blade angles
bcan along meridian length L
mBchange.Particularly, Fig. 7 shows corresponding to the meridian blade angles between hub curve 108 and guard shield curve 110
bmultiple curves.It should be noted that each curve between hub curve 108 and guard shield curve 110 can be corresponding with the mesosphere between the hub face 80 of blade 66 and guard shield face 82.As can be seen from the contrast of multiple curve, the meridian blade angles at hub face 80 place
busually the meridian blade angles at guard shield face 82 place can be greater than
b(that is, blade 66 can be more vertical at hub face 80 place).In addition, the meridian blade angles at two faces place
breach maximum in leading edge 86 and reach minimum at trailing edge 84.That is, meridian blade angles
breduce from leading edge 86 to trailing edge 84 generally.Still as shown in Figure 7, the meridian blade angles at leading edge 86 place
bcan change between about-5 ° to 30 °, and the meridian blade angles at trailing edge 84 place
bcan change between about-40 ° to-80 °.This blade angle distribution can contribute to reducing aerodynamic losses, and therefore improves performance and the efficiency of turbosupercharger 36.
As shown in Figure 8, the thickness T of blade 66
balso can along its meridian length L
mBchange.Especially, Fig. 8 shows corresponding to the meridian length L of blade 66 between hub face 80 and guard shield face 82 relative to blade 66
mBthickness T
bmultiple curves.Can find out from multiple curve, the thickness of blade 66 is at meridian length L
mBabout 60-80% place can reach maximum ga(u)ge T
bmax, about 10 millimeters, and be the thinnest at trailing edge 84 and leading edge 86 place.In one embodiment, maximum ga(u)ge T
bmaxcan at meridian length L
mBabout 68% place.Still as shown in Figure 8, the thickness along hub face 80 of blade 66 can much larger than the thickness along guard shield face 82.Finally, can be about 0.38 × T at the maximum ga(u)ge at leading edge 86 place
bmax, and can be about 0.61 × T at the maximum ga(u)ge at trailing edge 84 place
bmax.This level and smooth Bezier curve thickness distribution of blade 66 can improve the manufacturability of blade, and especially use side milling technique, its comparable alternative manufacturing process on cost is low.
Get back to Fig. 3, the exemplary geometric structure of the fin 76 of nozzle ring 62 of the present invention will be discussed now.For the purposes of the present invention, meridian length L
mVcan refer between the trailing edge 96 of fin 76 and leading edge 98 along the distance through the curved line of the longitudinal center of fin.Similar with blade 66, fin 76 also can along its curved in length, the corresponding meridian fin angle β of each formation
v, it is defined by following equation:
θ=angular coordinates, polar angle or cornerite
Z
m=along the local meridian coordinate of meridian length
R=local radial position
β
vmeridian fin angle ,=local
Thickness T
vthe distance of the curved line being approximately perpendicular to fin 76 between high pressure side 100 and low voltage side 102 can be referred to.Chord length L
cVthe crow flight distance between the trailing edge 96 of fin 76 and leading edge 98 can be referred to.Interval S
vcrow flight distance between the adjacent trailing edge 96 that can refer to adjacent fins 76.Solidity ratio SR
vchord length L can be defined as
cVand interval S
vvalue (SR
v=L
cV/ S
v).Width W
vthe distance between leading edge 86 place hub face 92 and guard shield face 94 can be referred to.Width compares WR
vwidth W can be defined as
vwith chord length L
cVratio (WR
v=W
v/ L
cV).Blade inlet guard shield tip radius r
1the distance of leading edge 86 in guard shield face 82 from the center of turbine wheel 48 to blade 66 can be referred to.Fin leading-edge radius r
2the distance of leading edge 98 in guard shield face 82 from the center of turbine wheel 48 to fin 76 can be referred to.Fin inlet radius compares IR
vfin leading-edge radius r can be defined as
2with blade inlet guard shield tip radius r
1ratio.Nozzle entrance established angle φ
vchord length L can be referred to
cVwith fin leading-edge radius r
2between angle.Fin trailing edge radius r
3the distance of trailing edge 96 in guard shield face 82 from the center of turbine wheel 48 to fin 76 can be referred to.Fin exit radius compares ER
vfin trailing edge radius r can be defined as
3with blade inlet guard shield tip radius r
1ratio.
Similar with blade 66, the geometrical construction of fin 76 of the present invention is selected as the Aerodynamic Flows angle providing expectation, and wherein the outlet port of nozzle ring 62 has the flow uniformity of improvement, increase the structural integrity of fin, and fin 76 has low torque load.Such as, each fin 76 can have the solidity ratio SR of about 0.7-1.2
v, around nozzle ring 62, comprise 13 to 25 fins 76.In one embodiment, the surrounding of nozzle ring 62 comprises 23 blades, and solidity ratio is about 1.11.The width of fin 76 compares WR
vcan be about 0.2-0.40.In one embodiment, this width compares WR
vfor about 0.23.The blade inlet radius ratio IR of fin 76 can be about 1.3-1.5.In one embodiment, fin inlet radius is about 1.36 than IR.The fin exit radius of fin 76 can be about 1.05-1.3 than ER.In one embodiment, fin inlet radius is about 1.19 than ER.Finally, the nozzle entrance established angle φ of fin 76
vit can be about 60 °-80 °.In one embodiment, nozzle entrance established angle φ
vfor about 74 °.Each these geometric properties can contribute to reducing aerodynamic losses, reduce the torque load(ing) of fin, and improve the structural integrity of fin 76, simultaneously with respect to being conducive to the smoothed curve improving manufacturability.
In addition, similar with blade 66, the meridian blade angles of fin 76
vcan along meridian length L
mVchange.Particularly, Fig. 9 and Figure 10 shows the curve of two kinds of different mode of executions of nozzle ring 62, and it corresponds to the meridian fin angle β from leading edge 98 to trailing edge 96
v.In often kind of these mode of execution, meridian fin angle β
vcan change in the scope of about 50-80 ° along its meridian length.
In the first mode of execution shown in Fig. 9, meridian fin angle β
vcan be different substantially along hub face 92 and guard shield face 94.Therefore, two curves separated are shown.Hub curve 112 can correspond to the meridian fin angle β along hub face 92
v, and guard shield curve 114 can correspond to the meridian fin angle β along guard shield face 94
v.In this embodiment, two curves 112 and 114 can share generally'S '-shaped curve along meridian length, show the shape in the chamber of fin 76.But along each some place on the meridian line between leading edge 98 and trailing edge 96, hub curve 112 can be greater than guard shield curve 114 substantially.
In the second mode of execution in Fig. 10, meridian fin angle β
vcan be equal substantially along hub face 92 and guard shield face 94.Therefore, a curve is merely illustrated.Figure 10 shows and corresponds to meridian fin angle β for both hub face 92 and guard shield face 94
vcurve 116.In this embodiment, fin 76 also can have the arc of generally'S '-shaped.In addition, in the present embodiment, the inclination angle from leading edge 98 to trailing edge 96 can be had, be shown as trend curve 118 here.Two kinds of mode of executions of above-mentioned fin 76 can use together with nozzle ring 62 according to the application expected.Hub face 92 and guard shield face 94 have two points of other fin angle distributions contribute to improving turbine bucket 66 vibratory response characteristic by reducing high cycle fatigue strain under wide in range operational condition.Hub face 92 and guard shield face 94 all have single fin angle distribution and can be more suitable for improving level aerodynamic quality under wide in range operational condition.
As shown in figure 11, the thickness T of fin 76
vcan along its meridian length L
mVchange.Especially, Figure 11 shows and the meridian length L of fin 76 relative to fin 76
mVthickness T
vcorresponding curve 120.As finding out from curve, the thickness of fin 76 does not change between hub face 92 and guard shield face 94.Thickness is at meridian length L
mVabout 20-50% place can reach the maximum ga(u)ge T of about 5.5 millimeters
vmax, and the thinnest at trailing edge 96 place.In one embodiment, the thickness of blade 66 is at meridian length L
mVabout 32% place reach maximum.Finally, can be about 0.25 × T at the maximum ga(u)ge at leading edge place
vmax, and can be about 0.09 × T at the maximum ga(u)ge at trailing edge place
vmax.In the mode similar with blade 66, this level and smooth thickness distribution of fin 76 can improve the manufacturability of fin.
industrial applicibility
Turbosupercharger of the present invention may be implemented in any employing in the power system application of super charge.Especially, the concrete geometrical construction of blade 66 and fin 76, angle, blade/vane face and thickness distribution can bring up the overall lower aerodynamic loss of turbo machine 32, and the performance therefore improved and efficiency.From nozzle ring 62 leave evenly and the flowing of good guidance can bring nozzle ring 62 and turbine wheel 48 evenly loading.This can contribute to reducing the cyclic loading on turbine wheel 48, extends the working life of turbine wheel 48.Due to exhaust stream can be substantially uniformly and be directed to each blade 66 well, the machinery that exhaust stream and turbine bucket geometrical construction due to misalignment cause and vibration loss can significantly reduce.In addition, nozzle ring 62 can have low degree of compaction with turbine wheel 48 compared with the axial turbine level of equivalence, therefore, can have less fin and blade.The minimizing of fin and blade can be equal to the reduction of manufacture cost.Finally, the level and smooth angle of blade 66 and fin 76 and thickness distribution can allow to use these assemblies of Flank machining, and it can be the alternative more cheap than other manufacturing process.
Can carry out much remodeling and variation for turbosupercharger of the present invention, this will be readily apparent to persons skilled in the art.Those skilled in the art passes through the specification of turbosupercharger of the present invention and the thinking of practice, and other mode of execution will be apparent.Specification and mode of execution should be considered to be only that exemplary, real scope is pointed out by following claim and their equivalent.
Claims (10)
1. a turbosupercharger, comprising:
Housing, it limits compressor protective cover and turbomachine shroud at least in part;
Compressor impeller, it is arranged in compressor protective cover;
Axle, it is connected to compressor impeller;
Turbine wheel, it to be arranged in turbomachine shroud and to be connected to one end relative with compressor impeller of axle, and turbine wheel comprises:
Be roughly the hub of annular; And
Multiple blade, it is arranged around annular hub radial direction, described multiple blade includes aerofoil, this aerofoil have be connected to annular hub hub face, relative with hub face and be oriented to towards the guard shield face of turbomachine shroud, trailing edge and the leading edge relative with trailing edge, the angle wherein between the pedestal of annular hub and leading edge is about 25-55 degree; And
Nozzle ring, this nozzle ring comprises:
Ring-type be roughly flat dish, it is positioned at the periphery of turbine wheel; And
Multiple fin, its radial direction of upper surface around dish is arranged, and wherein the arc of multiple fin described in each is cardinal principle S shape along its meridian length (LMV) from the leading edge of blade multiple described in each to trailing edge.
2. turbosupercharger according to claim 1, wherein, the solidity ratio of multiple blade described in each changes from about 0.8 to 1.2 between hub face and guard shield face.
3. turbosupercharger according to claim 1, wherein, the blade angle of multiple blade described in each reduces from hub face to guard shield face.
4. turbosupercharger according to claim 3, wherein, blade angle is maximum at leading edge place.
5. turbosupercharger according to claim 3, wherein, blade angle is minimum at trailing edge place.
6. turbosupercharger according to claim 1, wherein, the thickness of multiple blade described in each reduces from hub face to guard shield face, and the about 60-80% place of the meridian length of multiple blade is maximum described in each.
7. turbosupercharger according to claim 1, wherein, the leading edge of arc substantially along its meridian length from fin multiple described in each of the S shape of multiple fin described in each tilts to trailing edge.
8. turbosupercharger according to claim 1, wherein:
Multiple fin described in each comprises hub face and guard shield face; And
Fin angle along hub face is substantially different from the fin angle along guard shield face.
9. turbosupercharger according to claim 1, wherein:
Multiple fin described in each comprises hub face and guard shield face; And
Fin angle along hub face is substantially equal to the fin angle along guard shield face.
10. turbosupercharger according to claim 1, wherein, the thickness of multiple fin described in each is along its meridian length change, and this thickness is maximum at the about 20-50% place of meridian length.
Applications Claiming Priority (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US14/037,836 US20150086396A1 (en) | 2013-09-26 | 2013-09-26 | Turbocharger with mixed flow turbine stage |
| US14/037,836 | 2013-09-26 |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN104514631A true CN104514631A (en) | 2015-04-15 |
Family
ID=52691111
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201410494389.XA Pending CN104514631A (en) | 2013-09-26 | 2014-09-24 | Turbocharger with mixed flow turbine stage |
Country Status (2)
| Country | Link |
|---|---|
| US (1) | US20150086396A1 (en) |
| CN (1) | CN104514631A (en) |
Cited By (5)
| Publication number | Priority date | Publication date | Assignee | Title |
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| CN107304708A (en) * | 2016-04-19 | 2017-10-31 | 本田技研工业株式会社 | Turbomachinery |
| CN111042870A (en) * | 2018-10-11 | 2020-04-21 | 博格华纳公司 | Turbine wheel |
| CN111535872A (en) * | 2020-04-07 | 2020-08-14 | 东方电气集团东方汽轮机有限公司 | Bladeless transition mixed flow turbine structure |
| CN112434388A (en) * | 2020-11-20 | 2021-03-02 | 北京航天动力研究所 | Design method of high-pressure-ratio multistage axial flow type hydraulic turbine |
| CN114320477A (en) * | 2022-01-06 | 2022-04-12 | 上海万泽精密铸造有限公司 | A multistage impeller structure and turbo charger for turbo charger |
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| JPWO2015002066A1 (en) * | 2013-07-04 | 2017-02-23 | 株式会社Ihi | Compressor impeller, centrifugal compressor, compressor impeller processing method, and compressor impeller processing apparatus |
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| US9810238B2 (en) * | 2015-03-09 | 2017-11-07 | Caterpillar Inc. | Turbocharger with turbine shroud |
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| GB2555567A (en) * | 2016-09-21 | 2018-05-09 | Cummins Ltd | Turbine wheel for a turbo-machine |
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| EP3364039A1 (en) * | 2017-02-21 | 2018-08-22 | Siemens Aktiengesellschaft | Recirculation stage |
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| JP7140030B2 (en) * | 2019-03-28 | 2022-09-21 | 株式会社豊田自動織機 | Centrifugal compressor for fuel cell |
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| EP1904730B1 (en) * | 2005-07-19 | 2008-12-10 | Honeywell International Inc. | Variable nozzle turbocharger |
| US20120023936A1 (en) * | 2010-07-30 | 2012-02-02 | Caterpillar Inc. | Nozzled turbocharger turbine |
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| Publication number | Priority date | Publication date | Assignee | Title |
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| CN107304708A (en) * | 2016-04-19 | 2017-10-31 | 本田技研工业株式会社 | Turbomachinery |
| CN107304708B (en) * | 2016-04-19 | 2019-09-17 | 本田技研工业株式会社 | Turbomachinery |
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| CN111042870B (en) * | 2018-10-11 | 2023-09-12 | 博格华纳公司 | Turbine wheel |
| CN111535872A (en) * | 2020-04-07 | 2020-08-14 | 东方电气集团东方汽轮机有限公司 | Bladeless transition mixed flow turbine structure |
| CN111535872B (en) * | 2020-04-07 | 2022-01-11 | 东方电气集团东方汽轮机有限公司 | Bladeless transition mixed flow turbine structure |
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