WO2022263823A1 - Bague de buse pour une turbine à géométrie variable - Google Patents

Bague de buse pour une turbine à géométrie variable Download PDF

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Publication number
WO2022263823A1
WO2022263823A1 PCT/GB2022/051510 GB2022051510W WO2022263823A1 WO 2022263823 A1 WO2022263823 A1 WO 2022263823A1 GB 2022051510 W GB2022051510 W GB 2022051510W WO 2022263823 A1 WO2022263823 A1 WO 2022263823A1
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WO
WIPO (PCT)
Prior art keywords
nozzle
nozzle ring
intercept
corner
deck
Prior art date
Application number
PCT/GB2022/051510
Other languages
English (en)
Inventor
Stephen David HUGHES
George E. SANDFORD
Christopher J. Shaw
Original Assignee
Cummins Ltd
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Cummins Ltd filed Critical Cummins Ltd
Priority to GBGB2400649.6A priority Critical patent/GB202400649D0/en
Publication of WO2022263823A1 publication Critical patent/WO2022263823A1/fr

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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
    • 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
    • F01D17/141—Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits by means of shiftable members or valves obturating part of the flow path
    • F01D17/143—Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits by means of shiftable members or valves obturating part of the flow path the shiftable member being a wall, or part thereof of a radial diffuser
    • 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
    • F01D17/167—Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits by means of nozzle vanes of vanes moving in translation
    • 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
    • F05D2250/71—Shape curved

Definitions

  • the present invention relates a nozzle ring of a variable geometry turbine, such as a turbocharger, to a turbine incorporating the turbine, and to a turbo-charger incorporating the turbine.
  • Turbochargers are well-known devices for supplying air to the intake of an internal combustion engine at pressures above atmospheric pressure (boost pressures).
  • a conventional turbocharger essentially comprises an exhaust gas driven turbine wheel mounted on a rotatable shaft within a turbine housing. Rotation of the turbine wheel rotates a compressor wheel mounted on the other end of the shaft within a compressor housing. The compressor wheel delivers compressed air to the inlet manifold of the engine, thereby increasing engine power.
  • the turbocharger shaft is conventionally supported by journal and thrust bearings, including appropriate lubricating systems, located within a central bearing housing connected between the turbine and compressor wheel housing.
  • the turbine stage comprises a turbine chamber within which the turbine wheel is mounted; an annular inlet passageway defined between facing radial walls arranged around the turbine chamber; an inlet arranged around the inlet passageway; and an outlet passageway extending axially from the turbine chamber.
  • the passageways and chambers communicate such that pressurised exhaust gas admitted to the inlet chamber flows through the inlet passageway to the outlet passageway via the turbine and rotates the turbine wheel.
  • vanes referred to as nozzle vanes
  • Each vane is generally laminar, and is positioned with one radially outer surface arranged to oppose the motion of the exhaust gas within the inlet passageway, i.e. the circumferential component of the motion of the exhaust gas in the inlet passageway is such as to direct the exhaust gas against the outer surface of the vane.
  • Turbines may be of a fixed or variable geometry type.
  • Variable geometry type turbines differ from fixed geometry turbines in that the geometry of the inlet passageway can be varied to optimise gas flow velocities over a range of mass flow rates so that the power output of the turbine can be varied to suit varying engine demands.
  • a nozzle ring In one form of a variable geometry turbocharger, a nozzle ring carries a plurality of axially extending vanes, which extend into the air inlet, and through respective apertures (“slots”) in a shroud which forms a radially-extending wall of the air inlet.
  • the nozzle ring is axially movable by an actuator to control the width of the air passage. Movement of the nozzle ring also controls the degree to which the vanes project through the respective slots.
  • variable geometry turbocharger is shown in Figs. 1(a) and 1(b), taken from US 8,172,516, though with adapted annotation.
  • the illustrated variable geometry turbine comprises a turbine housing 1 defining an inlet chamber 2 to which gas from an internal combustion engine (not shown) is delivered.
  • the exhaust gas flows from the inlet chamber 2 to an outlet passageway 3 via an annular inlet passageway 4.
  • the inlet passageway 4 is defined on one side by the face of a movable annular wall member 5 which constitutes the nozzle ring, and on the opposite side by an annular shroud 6, which covers the opening of an annular recess 8 in the facing wall.
  • Fig. 1(c) illustrates three planes 101 which include the axis 100. These planes are referred to here as “radial planes”.
  • Fig. 1(b) is a cross-section of part of the turbine housing 1 using one of the radial planes 101 as the basis for the cross- section. Such a cross-section is referred to in this document as an “axial cross-section”.
  • Rotation of the compressor wheel 11 about the rotational axis 100 pressurizes ambient air present in an air inlet 12 and delivers the pressurized air to an air outlet 13 from which it is fed to an internal combustion engine (not shown).
  • the speed of the turbine wheel 9 is dependent upon the velocity of the gas passing through the annular inlet passageway 4.
  • the gas velocity is a function of the width of the inlet passageway 4, the width being adjustable by controlling the axial position of the nozzle ring 5.
  • the width of the passageway can be reduced by moving the nozzle ring “forward” (that is, from the left to the right in Fig. 1(a); the axial direction from the bearing assembly 14 towards the turbine wheel 9).
  • FIG. 1(a) shows the annular inlet passageway 4 closed down to a minimum width, whereas in Fig. 1(b) the inlet passageway 4 is shown fully open.
  • the nozzle ring 5 supports an array of circumferentially and equally spaced vanes 7, each of which extends across the inlet passageway 4.
  • the vanes 7 are orientated to deflect gas flowing through the inlet passageway 4 towards the direction of rotation of the turbine wheel 9.
  • the vanes 7 project through suitably configured slots in the shroud 6 and into the recess 8.
  • Each vane has an “inner” major surface which is closer to the rotational axis, and an “outer” major surface which is further away.
  • a pneumatically or hydraulically operated actuator 16 is operable to control the position of the nozzle ring 5 within an annular cavity 19 defined by a portion 26 of the turbine housing via an actuator output shaft (not shown), which is linked to a stirrup member (not shown).
  • the stirrup member in turn engages axially extending guide rods (not shown) that support the nozzle ring 5. Accordingly, by appropriate control of the actuator 16 the axial position of the guide rods and thus of the nozzle ring 5 can be controlled.
  • electrically operated actuators could be used in place of a pneumatically or hydraulically operated actuator.
  • the nozzle ring 5 has a generally annular portion 28 defining a substantially flat front (i.e. facing the shroud 6) surface 28 (the “nozzle deck”). Note that in some arrangements it is known for a portion of the front surface of the nozzle ring 5 to be a flat surface which is closer to the shroud 6 than a flat front surface of the nozzle ring which is radially inward of the vanes, and in this case the radially-inward flat surface is regarded as the nozzle deck 28.
  • the nozzle ring 5 further includes inner and outer annular flanges 17 and 18 which are substantially cylindrical about the axis 100.
  • the flanges 17, 18 extend axially rearward from the nozzle deck 28 into the annular cavity 19, and the cylindrical inner surface 29 of the flange 17 constitutes a cylindrical inner surface 29 of the nozzle ring.
  • the flat nozzle deck 28 and the cylindrical inner surface 29 of the nozzle ring meet in a rounded corner 30.
  • this corner has a radius of curvature (as measured in a cross- section using a plane 101 including the axis 100; that is, an axial cross-section) of 1.5mm.
  • the annular cavity 19 is separated by a wall 27 from a chamber 15.
  • Inner and outer sealing rings 20 and 21 are provided to seal the nozzle ring 5 with respect to inner and outer cylindrical surfaces of the annular cavity 19, while allowing the nozzle ring 5 to slide within the annular cavity 19.
  • the inner sealing ring 20 is supported within an annular groove 22 formed in the wall 27 facing into in the cavity 19, and bears against the inner annular flange 17 of the nozzle ring 5.
  • the outer sealing ring 21 is supported within an annular groove 23 in the annular flange 18 of the nozzle ring 5 and bears against the radially outermost internal surface of the cavity 19.
  • a first set of pressure balance apertures 25 is provided in the nozzle ring 5 within the vane passage defined between adjacent apertures, while a second set of pressure balance apertures 24 are provided in the nozzle ring 5 outside the radius of the nozzle vane passage.
  • the present invention aims to provide new and useful vane assemblies for use in a turbo machine, as well as new and useful turbo-machines (especially turbo-chargers) incorporating the vane assemblies.
  • the invention is based on the realisation that when the nozzle deck of a variable geometry turbine is spaced from the shroud (i.e. the inlet passage way is partially open), the shape of the corner in the nozzle ring proximate the radially inner edge of the nozzle deck is critical in determining whether the gas flow through the gas inlet is turbulent. In particular, making this corner gentler is associated with a reduction in turbulence at the turbine wheel, and a consequent improvement in efficiency.
  • the corner can be considered as extending between a “nozzle deck intercept” with the flat nozzle deck, and an “inner surface intersect” with the inner surface of the ring.
  • the present invention proposes that at at least one point on the corner (as viewed in an axial cross-section), intermediate the nozzle deck intercept and the inner surface intercept (and therefore spaced from both), the radius of curvature, R, of the corner is at least 2mm.
  • the concept may alternatively be defined in terms of a dimensionless quantity, namely the ratio of (i) the radius of curvature R at the at least one point, to (ii) a radial distance from the inner surface of the nozzle ring to the trailing edge of the vanes (which may for example be expressed as the radially-innermost radial position at which the fillet of the vane ends and the axially-extending vane surface begins).
  • it may be at least 0.3, at least 0.4, at least 0.5, at least 0.7, at least 1.0, or even at least 1.5 or 2.0.
  • Making the corner gentler typically increases the radial distance from the inner diameter of the nozzle ring to the beginning of the nozzle deck (“the nozzle deck intercept”). In known systems this distance is typically 1.5mm, but in embodiments of the present invention it is typically greater, for example at least 2mm, at least 2.5mm, at least 3mm, or even at least 4mm.
  • the corner is shaped such that its profile (when viewed in an axial cross- section) is a quarter of a circle having a diameter of 1 5mm, with the inner surface of the nozzle ring and the nozzle deck being respective tangents to the circle.
  • Simply increasing the radius of curvature of the corner would increase the axial distance d from inner surface intercept to the nozzle deck. If, when the nozzle ring is retracted to open the passage, the inner surface intercept encounters the seal on the inner surface of the nozzle ring, the seal is made less effective. Though this could be avoided by limiting the axial distance through which the nozzle ring is allowed to travel, that would reduce the degree to which the turbine can be controlled.
  • the invention proposes that either the corner is not formed with a circular cross-section, or that if it is, the inner surface of the nozzle ring is not a tangent to the circle.
  • the corner may be formed with a cross-section which is a portion of a circle which intercepts the inner surface of the nozzle ring non-tangentially with an obtuse internal angle.
  • the surfaces to either side of the inner surface intercept have different normal directions, such that the normal direction to the surface of the nozzle ring changes discontinuously at the inner surface intercept (this is referred to below as having an “edge break”).
  • the dimensionless “mismatch ratio” as the ratio of (i) the radius of curvature R, to (ii) the axial distance d from the inner surface intercept to the nozzle deck. Whereas in a conventional system, this ratio is equal to 1, in an embodiment it is preferably at least 1.4, and preferably at least 2.
  • the corner may be formed with a cross-section which is a portion of an ellipse.
  • the long axis of the ellipse may be substantially radial (i.e. normal to the rotational axis), and the short axis may be substantially axial (i.e. parallel to the rotational axis).
  • the cross-section of the corner may be a quarter of the ellipse.
  • the axial distance from the inner surface intercept to the nozzle deck (substantially half the short diameter of the ellipse) is less than the radial distance from the inner surface to the nozzle deck intercept (substantially half the long diameter of the ellipse).
  • the ratio between the long and short axes of the ellipse may be at least 1.5, and more preferably at least 1.8, at least 2 or even at least 2.5. Note that the maximum radius of curvature of the corner is very close to the nozzle deck intercept.
  • a “divergence ratio” This is defined as (i) the radial distance h from the inner surface of the nozzle ring to a point on the corner which is halfway in the axial direction from the inner surface intercept to the nozzle deck intercept, to (ii) the axial distance (d) from the inner surface intercept to the nozzle deck intercept.
  • Fig. 1 is composed of Fig. 1(a) which is an axial cross-section of a known variable geometry turbine, Fig. 1(b) which is a cross-section of a part of the turbine of Fig. 1(a), and Fig. 1(c) which illustrates radial planes which are used for form the cross-section of Fig. 1(b);
  • Fig. 2 shows a portion of a known nozzle ring
  • Fig. 3 is composed of Figs. 3(a)-3(c) which show portions of four respective nozzle rings which are embodiments of the invention;
  • Fig. 4 is composed of Figs. 4(a) and 4(b) which respectively show how the turbine stage efficiency depends upon the variable geometry opening for the comparative example of Fig. 2 and three realisations of the embodiment of Fig. 3(a), respectively for full load (Fig. 4(a)) and another operating point (Fig. 4(b));
  • Fig. 5 shows a portion of a further nozzle ring which is an embodiment of the invention
  • Fig. 6 shows the corners of Fig. 2, Fig. 3(c) and Fig. 5 overlaid, to explain the concept of a divergence ratio
  • Fig. 7 is composed of Figs. 7(a) and 7(b) which respectively show the turbine stage isentropic efficiency of the comparative example of Fig. 2 and four realisations of the embodiments of Figs. 3(c) and 5 for two different expansion ratios, respectively with a 3mm variable geometry opening (Fig. 7(a)) and a 5.5mm variable geometry opening (Fig. 7(b)); and
  • Figs 8(a) and (b) are introduced to demonstrate the value of the divergence ratio for the corners of Fig. 2, Fig. 3(c) and Fig. 5.
  • a portion of a known nozzle ring 5 is shown, in a cross-sectional view which is the intersection between the portion of the nozzle ring 5 with a plane 101 including the axis 100 as described above with reference to Fig. 1(c). That is, the plane lies in a single angular position about the axis 100, and Fig. 2 is an “axial cross-section”.
  • the nozzle ring 5 of Fig. 2 may be the nozzle ring 5 shown in Figs. 1(a) and 1(b).
  • the nozzle ring 5 has a substantially flat nozzle deck 28, which faces forwards towards the inlet passageway 4 (to the right in Fig. 2). From the nozzle deck 28 vanes 7 project forward. Gas travels through the inlet passageway 4 in a generally radially inward direction (i.e. the down direction in Fig. 2).
  • An actuator can reciprocate the nozzle ring 5 in the forward-backward (i.e. right-left) direction.
  • Fig. 2 shows the nozzle ring 5 in the furthest rearward position in its range of possible positions.
  • the inlet passageway 4 is axially widest, maximising the flow of gas to the turbine wheel.
  • the nozzle deck 28 bears against a shroud (not shown) on the other (right) side of the inlet passageway 4, and the vanes 7 pass through slots in the shroud.
  • the inlet passageway 4 which is defined between the nozzle deck 28 and the shroud, is closed, and gas cannot pass to the turbine wheel.
  • An annular inner sealing ring 20 in the annular groove 22 in the turbine housing 27 seals the gap between the turbine housing 27 and the inner surface 31 of the nozzle ring 5.
  • the known nozzle ring 5 includes inner and outer flanges
  • the portion of the nozzle ring 5 shown in Fig. 2 only includes the inner flange 17, and for simplicity only its cylindrical inner surface 31 is shown in Fig. 2.
  • surface 31 is the inner surface of the nozzle ring 5.
  • Each vane 7 typically meets the nozzle deck 28 at filleted surfaces 34, 35.
  • the intercept of the inner filleted surface 35 with an axially-extending line 41 on the vane 7 is denoted 37.
  • point 35 is a transition between the portion 35 of the vane which is concave in this cross-section, and the straight line 41.
  • the vane 7 and the fillets 34, 35 appear different according to which plane including the axis 100 is used to form the cross-section (i.e. according to which angular position about the axis 100 Fig. 2 represents).
  • the plane which is used to form Fig. 2 is the one for which the point 37 is closest to the axis.
  • straight line 41 is the trailing edge of the vane 7, i.e. the axially-extending line on the axially-extending portion of the vane surface which is closest to the rotational axis (furthest in the direction in which the gas travels).
  • An axial line which is a continuation of the trailing edge 41 is labelled 38.
  • the radial distance from the inner surface 31 of the nozzle ring 5 to the line 38 is denoted as distance q. Since the trailing edge 41 extends axially, all points along it have the same radial distance q from the inner surface 31 of the nozzle ring.
  • distance q denotes the radial distance from the inner surface 31 of the nozzle ring 5 to the trailing edge 41 of the vanes 7 (this is the same for all the vanes 7, since the vanes 7 have rotational symmetry with each other about the rotational axis). In a known system, distance q may for example be about 5.5mm.
  • Portions of the nozzle ring 5 which are radially inward of the radially-inner fillet 35 (i.e. below fillet 35 as seen in Fig. 2) are rotationally symmetric about the axis 100. That is, they appear the same irrespective of which plane including the axis 100 is used to form the cross-section.
  • the nozzle deck 28 and the inner surface 31 of the nozzle ring are joined by a corner 30 which in cross-section is a quarter of a circle having a radius denoted R and a centre 40.
  • a radial line 43 from the centre 40 intercepts the surface of the nozzle ring 5 at the point where the inner surface 31 meets the corner 30. This point is referred to as the inner surface intercept 47.
  • An axial line 42 from the centre 40 intercepts the surface of the nozzle deck 28 at the point where the nozzle deck 28 meets the corner 30. This point is referred to as the nozzle deck intercept 46.
  • the radial distance from the inner surface 31 of the nozzle ring 5 to the nozzle deck intercept 46 is denoted p.
  • distance p is equal to the radius R, and is conventionally 1.5mm.
  • a portion is shown of a nozzle ring 51 which is an embodiment of the invention.
  • the corner 39 has a larger radius R (and thus, a larger value of distance d) than of the corner 30 of Fig. 2.
  • the radius R may be at least 3mm.
  • the inner surface 31 of the nozzle ring, and the nozzle deck 28, are tangents to the corner 30.
  • the distance q does not differ between the comparative example of Fig. 2 and the embodiment of Fig.
  • portions of the nozzle ring 51 which are radially inward of the radially-inner fillet are rotationally symmetric about the axis 100.
  • the inner sealing ring 20 has a forward-facing surface 201 and a rearward-facing surface 202. As shown in Fig. 3, the inner surface intercept 47 is closer to the forward-facing surface 201 of the inner sealing ring 20 than in Fig. 2. This causes the risk that due to manufacturing tolerances, at the fully open position of the nozzle ring 5 the inner surface intercept 47 may be in axial register with the inner sealing ring 20 (i.e. the inner surface intercept 47 may be axially rearward of the forward-facing surface 201), which may cause the sealing to fail.
  • the inner sealing ring 20 narrower in the axial direction.
  • the rearward-facing surface 202 of the inner sealing ring 20 may be in the same position, but the forward-facing surface 201 of the inner sealing ring 20 could be at an axial position indicated by the dashed line 203.
  • an axially- narrower inner sealing ring 20 would be less effective in sealing.
  • Fig. 4(a) shows the relationship between the stage efficiency of the turbine (the vertical axis, as a percentage) to the width of the inlet passage 4 (in mm) for systems with a distance q of 9mm under “full load” conditions.
  • Lines 71, 72 and 73 respectively show the embodiment of Fig. 3(a) in realisations with p and R both equal to 3mm, 6mm and 9mm (i.e. an R/q ratio of .54, 1.1 and 1.64, and mismatch ratio again equal to 1).
  • Fig. 4(b) shows the relationship between the stage efficiency of the turbine (the vertical axis, as a percentage) to the width of the inlet passage 4 (in mm) for systems with a distance q of 9mm under “part load” conditions.
  • Lines 76, 77 and 78 respectively show the embodiment of Fig. 3(a) in realisations with R (and p) respectively equal to 3mm, 6mm and 9mm (i.e. an R/q ratio of .54, 1.1 and 1.64). Again, all these lines exhibit a higher efficiency than the comparative example of Fig.
  • Fig. 3(b) shows portion of a nozzle ring 52 which is a further embodiment of the invention.
  • the nozzle ring 52 is identical to the nozzle ring 51 of Fig. 3(a), except that the corner 3 is replaced by a frusto-conical surface 45 which is rotationally symmetric about the axis 100.
  • This frusto-conical surface 45 appears as a straight line in Fig. 3(b), and thus has an infinitely high radius of curvature at all points intermediate the nozzle deck intercept 46 and the inner surface intercept 47.
  • the radial position of the nozzle deck intercept 46 can be chosen to be further from the inner surface 31 of the nozzle ring 52 than in the known nozzle ring of Fig. 2.
  • the flow separation illustrated in Fig. 2 is reduced.
  • the nozzle ring 52 has two sharp corners (“edge breaks”) at the nozzle deck intercept 46 where the frusto- conical surface 45 intercepts the nozzle deck 28, and at the inner surface intercept 47 where the frusto-conical surface 45 intercepts the inner surface 31 of the nozzle ring 52. These edge-breaks may themselves lead to turbulence.
  • Fig. 3(c) shows portion of a nozzle ring 53 which is a further embodiment of the invention.
  • the nozzle ring 52 is identical to the nozzle ring 51 of Fig. 3(a), except that the corner 39 is replaced by a corner 50.
  • the corner 50 is a portion of a circle 48 having a centre 40, but the corner 50 subtends an angle of less than 90 degrees about the centre 40.
  • the radius R of the circle 48 can thus be greater than that of the corner 30 of the nozzle ring 5 of Fig. 2, even if the nozzle deck intercept 46 and the inner surface 47 are at substantially the same positions. Again, it has been found that by avoiding having a corner 30 as in the nozzle ring 5 of Fig. 2, flow separation can be avoided.
  • mismatch ratio the ratio of R/d.
  • Fig. 3(c) like that of Fig. 3(b) exhibits “edge breaks” (sharp angles) at the nozzle deck intercept 46 and the inner surface intercept 47, though they are not as large as in Fig. 3(b) for a given value of p.
  • FIG. 3(a) When the corner is viewed in an axial cross-section so that it appears as a part of a circle, the nozzle deck 28 appears as a line which is tangential to the circle (as in Fig. 3(a)), but the inner surface 31 of the nozzle ring appears as a line which is not a tangent to the circle (as in Fig.
  • FIG. 5 shows a portion of a nozzle ring 54 which is further embodiment of the invention and which avoids edge breaks, while still endangering the sealing between the nozzle ring to a lesser degree than the embodiment of Fig. 3(a).
  • the nozzle ring 54 of Fig. 5 is identical to the nozzle ring 5 of Fig. 2 and the nozzle ring 51 of Fig. 3(a) except that the corners 30, 39 are replaced by a corner 56 between the inner surface 31 of the nozzle ring 54 and the flat nozzle deck 28 which (like the corners 30, 39, 50) is rotationally symmetric about the axis 100, but which in an axial cross-section is one quarter of an ellipse 55.
  • the centre of the ellipse 55 is the point 40.
  • the long axis of the ellipse has a length A
  • the short axis of the ellipse has a length B. Since A can be chosen independently of B, B can be chosen to be small enough to provide good axial clearance between the inner surface intercept 47 and the inner seal ring 20, and A can be chosen to be sufficiently large that the corner 56 is gentle enough that flow separation (as shown in Fig.. 2) does not occur.
  • the maximum radius of curvature at any point on the corner intermediate the intercepts 46 and 47 is referred to as R max .
  • R max the maximum radius of curvature at any point on the corner intermediate the intercepts 46 and 47.
  • the point with the maximum radius of curvature is substantially at the nozzle deck intercept 46, and the radius of curvature is equal to A 2 /2B.
  • the radius of curvature is the same value R at all points on the corner, so R max is equal to R.
  • distance h for any of the corners 30, 39, 50, 56 as the radial distance from the inner surface to the point on the corner 30, 39, 50, 56 which intercepts the line 57.
  • Fig. 7(a) shows for each of these 5 realisations (the horizontal axis is the realisation number), the corresponding turbine stage isentropic efficiency as a percentage if the width of the inlet passage 4 is 3mm for a turbine expansion ratio of 1.4 (points 61) and 2.6 (points 62).
  • Fig. 7(b) shows for each of these 5 realisations (the horizontal axis is the realisation number), the corresponding turbine stage isentropic efficiency as a percentage if the width of the inlet passage 4 is 5.5mm for a turbine expansion ratio of 1.4 (points 63) and 2.6 (points 64).
  • Fig. 8(a) where line 80 represents one of the corners 30, 50 with a radius of curvature R. Define x as R/d.
  • the corner 50 of the embodiment of Fig. 3(c) corresponds to the case that x is greater than one.
  • the radial line from centre 40 to the inner surface 31 is denoted 81
  • its intercept with the inner surface 31 is denoted 88.
  • the angle between radial line 81 and the line from the centre 40 to the inner surface intercept 47, is denoted Q.
  • Point 86 is the point on the line 57 in Fig. 6 which is at the same radial position as the inner surface 31.
  • Point 87 the intercept of the line 57 with the corner 30, 50.
  • distance h is the distance from point 86 to point 87.
  • f denotes the angle between line 81 and a line 83 from centre 40 to the point 87.
  • Eqns. (2) and (3) determine h for any value of x and Q.
  • x 1
  • R 1.42 times distance d (as in realisation 2 of Table 1)
  • x 1.42
  • Q 15°
  • f 37.7°
  • h/d 0.248.
  • Q 24.295°
  • h/d 0.134A/B.

Abstract

La présente invention concerne une turbine pour une turbomachine, dans laquelle, au niveau d'une entrée de gaz, des aubes s'étendent à partir d'une bague de buse à travers des fentes dans une enveloppe. La bague de buse possède un coin radialement interne qui est une courbe plus douce que dans un système classique, et il a été constaté que cela conduit à un rendement amélioré de la turbine.
PCT/GB2022/051510 2021-06-17 2022-06-15 Bague de buse pour une turbine à géométrie variable WO2022263823A1 (fr)

Priority Applications (1)

Application Number Priority Date Filing Date Title
GBGB2400649.6A GB202400649D0 (en) 2021-06-17 2022-06-15 Nozzle ring for a variable geometry turbine

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
GB2108695.4 2021-06-17
GBGB2108695.4A GB202108695D0 (en) 2021-06-17 2021-06-17 Nozzle ring for a variable geometry turbine

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WO2022263823A1 true WO2022263823A1 (fr) 2022-12-22

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Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2006131724A1 (fr) * 2005-06-07 2006-12-14 Cummins Turbo Technologies Limited Turbine a geometrie variable
EP1900908A2 (fr) * 2006-09-12 2008-03-19 Iveco Motorenforschung AG Turbine à géométrie variable
WO2011015908A1 (fr) * 2009-08-04 2011-02-10 Renault Trucks Turbine à géométrie variable
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