Nozzle ring for a variable geometry turbine

Abstract

A turbine for a turbo-machine is proposed in which, at a gas inlet, vanes, extend from a nozzle ring though slots in a shroud. The nozzle ring has a radially-inner corner which is a gentler curve than

Description

Field of the invention 5 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. Background of the invention 10 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 15 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 20 wheel housing. In known turbochargers, 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 25 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. 30 It is known to improve turbine performance by providing vanes, referred to as nozzle vanes, in the inlet passageway so as to deflect gas flowing through the inlet passageway towards the direction of rotation of the turbine wheel. 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. 10 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. 15 20 An example of such a 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. 25 30 Gas flowing from the inlet chamber 2 to the outlet passageway 3 passes over a turbine wheel 9 and as a result torque is applied to a turbocharger shaft 10 supported by a bearing assembly 14 that drives a compressor wheel 11. The turbine is arranged to rotate about a rotational axis 100. 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 crosssection. 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. For a fixed rate of mass of gas flowing into the inlet passageway, 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). As the width of the inlet passageway 4 is reduced, the velocity of the gas passing through it increases. 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. 10 15 20 25 30 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. When the nozzle ring 5 is proximate to the annular shroud 6 and to the facing wall, 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. It will be appreciated that 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. Typically, this corner has a radius of curvature (as measured in a cross-5 section using a plane 101 including the axis 100; that is, an axial cross-section) of 1,5mm. 10 15 The annular cavity 19 is separated by a wall 27 from a chamber 15. Inner and outer sealing rings 20 and 21, respectively, 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. It will be appreciated that the inner sealing ring 20 could instead be mounted in an annular groove in the flange 17 rather than as shown, and / or that the outer sealing ring 21 could be mounted within an annular groove provided within the radially-outerwall of the cavity rather than as shown. 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. Summary of the invention 25 The present invention aims to provide new and useful vane assemblies for use in a turbomachine, 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 30 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. For example, 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. Whereas in a conventional system, the radius of curvature of the corner in a cross-section including the axis of the nozzle ring (an axial cross-section) is about 5 1,5mm, 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. 10 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). In a known system, the radial 15 distance from the inner surface to the trailing edge of the vanes may be about 5.5mm (e.g. 5.5175 mm), so the ratio is at least 2 / 5.5=0.36. In embodiments of the present concept, 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. 20 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. 25 Conventionally, the corner is shaped such that its profile (when viewed in an axial crosssection) 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 30 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. Alternatively, it would be possible to 35 position the seal axially further from the gas inlet, but this would require the inner flange to be longer so that the seal is still effective when the nozzle ring is in the advanced position. This in turn increases the required size of the annular cavity in which the inner flange is located. 5 For that reason, the invention proposes that the corner is formed with a circular crosssection, but the inner surface of the nozzle ring is not a tangent to the circle. This make it possible for the inner surface intercept to be closer axially to the nozzle deck than the radius of curvature of the corner at least one point, or even over its whole extent. 10 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. Thus, 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 15 break”). "d" C\J We may define 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 20 preferably at least 1.4, and preferably at least 2. Another way of defining the corner is using a “divergence ratio”. This is defined as (i) the 25 radial distance h from the inner surface of the nozzle ring to a point on the corner which is half way 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. 30 For a conventional corner, the divergence ratio is h / d=0.134. In certain embodiments of the invention, the value of the divergence ratio is higher, such as at least 0.2 or at least 0.3. Embodiments of the invention will now be described for the sake of example only, with reference to the following drawings in which: Fig. 1 is composed of Fig. 1(a) which is an axial cross-section of a known variable 5 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; 10 Fig. 3 is composed of Figs. 3(a)-3(c) which show portions of four respective nozzle rings, of which Fig. 3(c) is an embodiment of the invention; 15 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 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; 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 25 isentropic efficiency of the comparative example of Fig. 2 and four realisations of the embodiments of Figs. 3(c) and the nozzle ring of Fig. 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 30 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. In the figures, like elements are labelled by the same reference numerals. Referring to Fig. 2, 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”. 10 15 20 25 30 The nozzle ring 5 of Fig. 2 may be the nozzle ring 5 shown in Figs. 1(a) and 1(b). As mentioned above, 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 (not shown) 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. When the nozzle ring 5 is in this “fully open” position, the inlet passageway 4 is axially widest, maximising the flow of gas to the turbine wheel. In the furthest forward position of the nozzle ring 5 (the “closed” position), 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. Thus, 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. As the nozzle ring moves forward to the closed position from the position shown in Fig. 2, the cylindrical inner surface 31 of the flange 17 slides over the inner sealing ring 20, until the inner sealing ring 20 is axially in register with the area 32 proximate the rearward end of the flange 17. Although, as shown in Fig. 1(b), 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. Thus, 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. In other words, point 35 is a transition between the portion 35 of the vane which is concave in this cross-section, and the straight line 41. Note that the vane 7 and the fillets 5 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. In this case, 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 10 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. 15 Thus, distance q denotes the radial distance from the inner surface 31 of the nozzle ring C\J 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. CD 20 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 25 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 30 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. Conventionally, distance p is equal to the radius R, and is conventionally 1.5mm. Thus, the ratio of the distance p to distance q is 1.5mm / 5.5mm 35 =.27. This is also the ratio R / q. It has been discovered that the conventional nozzle ring 5 leads to disadvantageous gas flow. Considering gas flow from points indicated as a and b, the gas travels respectively to points c and f. Flow separation occurs leading to a low pressure position e which 5 receives substantially none of the exhaust gas. This in turn leads to a risk of turbulence in the region c. This effect reduces the efficiency of the turbine. 10 15 20 25 30 Turning to Fig. 3(a), a portion is shown of a nozzle ring 51 which is a comparative example. This differs from the comparative example of Fig. 2 only in that, in Fig. 3(a), the corner 39 has a larger radius R (and thus, a larger value of distance d) than of the corner 30 of Fig. 2. For example, the radius R may be at least 3mm. As in the known nozzle ring of Fig. 2, 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 examples of Fig. 2 and Fig. 3(a), so the ratio of the distance p, which is the same as the radius R, to the distance q is 3mm / 5.5mm=0.55. As in the comparative example of Fig. 2, portions of the nozzle ring 51 which are radially inward of the radially-inner fillet are rotationally symmetric about the axis 100. It has been observed that in this case, less turbulent flow is observed, with reduced flow separation. This increases the turbine efficiency. 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. One way of addressing this would be to make the inner sealing ring 20 narrower in the axial direction. For example, 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. However, an axially-narrower inner sealing ring 20 would be less effective in sealing. Another alternative would be to locate the annular groove 22 further rearward than in the nozzle rings 5, 51 of Figs. 2 and 3, so that the inner sealing ring 20 would be further rearward. However, doing this would require that the inner flange 17 is extended rearwardly, so that even when the nozzle ring is in its most forward position (the closed 5 position), the inner sealing ring 20 is in still register with a portion 32 of the inner surface 31. This is disadvantageous because it increases the required size of the turbine housing. The increase in the size of the turbine is itself undesirable, but it also implies more metal is needed, and thus a higher production cost and higher thermal inertia. 10 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. Line 70 shows the case of the comparative example of Fig. 2 with a value of R=1.5mm (i.e. a R / q ratio of .27 if q is 5.5mm). Lines 71, 72 and 73 respectively show the example of Fig. 3(a) in realisations with p and R 15 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). Note that all these lines exhibit a higher efficiency than the comparative example of Fig. 2, but the results are close to each other, showing that provided the R / q ratio is above a threshold amount (which may be about .4) the exact value of the R / q ratio is not very important because flow separation has been avoided. CD 20 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. Line 75 shows the case of the comparative example of Fig. 2 with a value of R=1.5mm (i.e. a R / q ratio of .27). Lines 25 76, 77 and 78 respectively show the example 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. 2, particularly for a lower width of the nozzle gap. Again, the results are close to each other, showing that provided the R / q ratio is above a threshold amount (which may be about 30 .4) the exact value of the R / q ratio is not very important because flow separation has been avoided. Fig. 3(b) shows portion of a nozzle ring 52 which is a further comparative example. The nozzle ring 52 is identical to the nozzle ring 51 of Fig. 3(a), except that the corner 3 is 35 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 5 known nozzle ring of Fig. 2. By avoiding having a corner 30 as in the nozzle ring 5 of Fig. 2, the flow separation illustrated in Fig. 2 is reduced. However, 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. 10 These edge-breaks may themselves lead to turbulence. Fig. 3(c) shows portion of a nozzle ring 53 which is an 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 15 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. CD 20 We define the axial distance between the inner surface intercept 47 and the nozzle deck 28 as the distance d. If centre 40 is equidistant from the inner surface 31 and the nozzle deck 28, then d=p. We define the “mismatch ratio” as the ratio of R / d. Thus, the mismatch ratio is equal to 1 for the known corner 30 of Fig. 2, and for the corner 39 of Fig. 3(a), but 25 greater than 1 for the corner 50 of Fig. 3(c). Note that the embodiment of Fig. 3(c), like the example 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. 30 Further comparative examples combine features of Fig. 3(a) and Fig. 3(b) by arranging for the corner, which as in those examples is portion of a circle as viewed in an axial cross-section, to meet one of the nozzle deck and the inner surface tangentially and the other of the nozzle deck and the inner surface non-tangentially. In other words, in a first 35 possibility, 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. 3(c)); instead there is an edge break (sharp corer) at the inner surface intercept 47. Conversely, in a second possibility, when the corner is viewed in an axial cross-section so that it appears as a part of a circle, the inner surface 31 of the nozzle ring appears as a line which is tangential to the circle (as in Fig. 3(a)), but the nozzle deck 28 of the nozzle ring appears as a line which is not a tangent to the circle (as in Fig. 3(c)); instead, there is an edge break at the nozzle deck intercept 46. 10 15 20 25 30 By contrast Fig. 5 shows a portion of a nozzle ring 54 which is further comparative example and which avoids edge breaks, while still endangering the sealing between the nozzle ring to a lesser degree than the example of Fig. 3(a). Specifically, 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. An axial line 42 though the centre 40 meets the nozzle deck 28 at the nozzle deck intercept 46, and a radial line 43 through the centre 40 meets the inner surface 31 at the inner surface intercept 46. The long axis of the ellipse has a length A, and 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. It is a well-known property of an ellipse that its radius of curvature is A2 / B at the end of its short axis (i.e. at the nozzle deck intercept 46) and B2 / 2A at the end of its long axis (i.e. at the inner surface intercept 47). Thus, provided A2 / 2B is greater than 2mm, there is at least one point on the corner 56 intermediate between the nozzle deck intercept 46 and the inner surface intercept 47, which has a radius of curvature of at least 2mm (i.e. very close to the nozzle deck intercept 46). For any of the embodiments, the maximum radius of curvature at any point on the corner intermediate the intercepts 46 and 47 is referred to as Rmax. In the case of the elliptical 29 10 24 corner 56, the point with the maximum radius of curvature is substantially at the nozzle deck intercept 46, and the radius of curvature is equal to A2 / 2B. For the circular corners 30, 39, 50, the radius of curvature is the same value R at all points on the corner, so Rmax is equal to R. 5 Another way of quantifying the concept of the nozzle ring having a gentle inner corner is by considering the rate at which, proximate the inner surface intercept 47, the corners 30, 39, 50, 56, diverge from the axial direction. In Fig. 6(a), the corners 30, 50, 56 are overlain assuming that the axial distance d from the inner surface intercept 47 to the 10 nozzle deck 28 is the same in each case. It will be seen that, moving axially (to the right) starting from the inner surface intercept 47, the corners 50, 56 diverge from the axial direction more rapidly than the corner 30 of the known nozzle ring. This can be quantified by considering an axial position halfway between the inner surface intercept 47 and the nozzle deck 28. This axial position is indicated by the line 57. We define distance h for 15 any of the corners 30, 39, 50, 56 as the radial distance from the inner surface to the point on the comer 30, 39, 50, 56 which intercepts the line 57. We define the “divergence ratio” for each of the covers 30, 39, 53, 56 as the value of h / d. For each of the curves 30, 53, 56 It can be shown (see Appendix 1) that h / d=0.134 for corners 30 and 39, h / d=0.322 for corner 50 when the radius R is twice d, and h / d=0.134A / B for corner 56. 20 To compare the efficiency of the comparative example of Fig. 2 and the embodiments of Fig. 3(c) and Fig. 5, various realisations were tested as shown in Table 1. Realisation number Shown in figure Numerical parameters Mismatch ratio Rmax / C| h / d 1 Fig. 2 q=5.5mm, R=p=1.5mm 1 0.27 0.134 2 Fig. 3(c) q=5.5mm, R=2.13mm, p=d=1.5mm 1.42 0.39 0.248 3 Fig. 3(c) q=5.5mm, R=3mm, p=d=1.5mm 2 0.55 0.322 4 Fig. 5 q=5.5mm, A=2p=9mm, B=2d=3mm 2.45 0.4 5 Fig. 5 q=5.5mm, A=2p=9mm, B=2d=4.25mm 1.73 0.283 Table 1 5 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). 10 15 Similarly, 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). It will be seen that all the four nozzle rings (realisations 2-5) are more efficient for both widths of the inlet passage 4 than the comparative example (realisation 1), and for both the expansion ratios. The most efficient realisation for the lower expansion ratio (1.4) is realisation 5, but at the higher expansion ratio (2.6) realisations 3 and 5 are roughly equally efficient. 20 Thus a preferred feature of the corner is that the value of h / d is at least 0.2, and preferably it is at least 0.25. Appendix 1 25 To find the value of h in Fig. 6 for the corner 30 of Fig. 2, and for the corner 50 of Fig. 3(c) assuming that p=d, consider Fig. 8(a), where line 80 represents one of the corners 30, 50 with a radius of curvature R. Define x as R / d. Thus, the known corner 30 corresponds to the case x=1, and the corner 50 of the embodiment of Fig. 3(c) 30 corresponds to the case that x is greater than one. For simplicity, the units of measurement in Fig. 8(a) are chosen so that d=1. The radial line from centre 40 to the inner surface 31 is denoted 81, and 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 6. 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. Thus, distance h is the distance from point 86 to point 87. 5 Since d and p have value 1, the line from the nozzle deck intercept 46 to the inner surface intercept 47 has distance V2, so half this distance is 1 / V2. It is apparent from Fig. 8(a) that 0 is given by sin (45°- 0)=l / xV2, or equivalently 0=(45° -sirT^l / xV^). (1) 10 Eqn. (1) determines 0 for any given x. For example, for x=1, 0 = 0 degrees (which is correct because in the case x=1, centre 40 is axially in register with the inner surface intercept 47). It is also apparent from Fig. 8(a) that the distance between points 86 and 88 is x sin 0 +1 / 2. 15 Turning to Fig. 8(b), the trapezium having points 40, 87, 88 and 86 is shown. <p denotes C\j the angle between line 81 and a line 83 from centre 40 to the point 87. From Fig. 8(b) xsin <p= x sin 0 +1 / 2, (2) and 20 h=x cos 0 - x cos (3) Eqns. (2) and (3) determine h for any value of x and 0. For example, from (1), (2) and (3), for corner 30 (i.e. the case x=1), 0 = 0, <p =30° and 25 h / d= 1-(^ / 3 / 2) = 0.134. For corner 50, in the case that R is 1.42 times distance d (as in realisation 2 of Table 1), so x=1.42, 0 = 15°, (p= 37.7° and h / d=0.248. For corner 50, in the case that R is twice distance d, x=2, 0 = 24.295°, <p =41.41°and h / d=0.322. Using the first of these results, for an ellipse with a ratio of a long radial axis and a short 30 axial axis equal to A / B, h / d is 0.134A / B.

Claims

1. A nozzle ring for a turbine, the nozzle ring having:a flat annular nozzle deck surface encircling an axis and facing in a direction 5 parallel to the axis;a cylindrical inner surface of the nozzle ring encircling the axis;a plurality of vanes extending axially from the nozzle deck;a corner surface which is rotationally symmetric about the axis and extends from an inner surface intercept on the cylindrical inner surface to a nozzle deck intercept on 10 the nozzle deck surface;a cross-section of the corner surface, measured in a plane including the axis, between the nozzle deck intercept and the inner surface intercept being a portion of a circle having a radius of curvature R greater than 2mm, and a mismatch ratio of (i) the radius of curvature R, to (ii) an axial distance d from the inner surface intercept to the 15 nozzle deck intercept is at least 1.4.C\J 2. A nozzle ring according to claim 1 in which the ratio of (i) the radius of curvatureR at the at least one point, to (ii) a radial distance q from the cylindrical inner surface of the nozzle ring to the trailing edge of the vanes is more than 0.3.O 20CM 3. A nozzle ring according to claim 1 or claim 2 in which the mismatch ratio is atleast 2.

4. A nozzle ring according to any preceding claim in which a divergence ratio which 25 is the ratio of (i) the radial distance h from the cylindrical inner surface of the nozzle ring to a point on the corner which is half way 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, is greater than 0.2.30 5. A nozzle ring according to claim 4 in which the divergence ratio is greater than0.3.

6. A turbine having:a turbine housing defining a gas inlet; and35 a turbine wheel positioned for rotation within the turbine housing about an axis;a ring-shaped shroud defining a plurality of slots, and encircling the axis;a nozzle ring according to any preceding claim; anda drive mechanism for selectively moving the nozzle ring towards or away from the shroud.5 7. A turbocharger comprising a turbine according to claim 6.29 10 24

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