WO2020229616A1

WO2020229616A1 - Variable geometry turbine

Info

Publication number: WO2020229616A1
Application number: PCT/EP2020/063505
Authority: WO
Inventors: Paris AMY; Fahim PATEL; James Mcewen; Christopher Parry; Andrew Sullivan; Paul GHOSH; Ivan Arbuckle
Original assignee: Cummins Ltd
Priority date: 2019-05-14
Filing date: 2020-05-14
Publication date: 2020-11-19
Also published as: GB2585634A; GB201906791D0

Abstract

A variable geometry turbine comprising a turbine wheel (5) supported for rotation about a turbine axis (4a); an annular inlet passageway (9) surrounding the turbine wheel (5) and defined between a first radial inlet surface (10) of a first wall member (11) and a second radial inlet surface of a second wall member (12), at least one of said first and second wall members (11, 12) being moveable along the turbine axis to vary the size of the inlet passageway (9); an array of vanes (14) extending across the inlet passageway (9), said vanes (14) being connected to said first wall member (11); a plurality of cavities (13) located in the second wall member (12) or to a rear side of the second wall member, wherein the rear side of the second wall member (12) is located axially beyond the second wall member (12) with respect to the first wall member (11); an array of vane slots (25) in the second radial inlet surface, complementary to said array of vanes (14), said vane slots being configured to receive said vanes and permit the vanes to travel into said plurality of cavities (13), to accommodate relative movement between the first and second wall members (11, 12); wherein the plurality of cavities (13) comprises a first cavity (41), and a second cavity (42) separated from the first cavity by a separating portion (48).

Description

VARIABLE GEOMETRY TURBINE

The present invention relates to a turbine. In particular, it relates to a variable geometry turbine having vanes that extend across an inlet passageway of 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 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 intake 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.

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 passageway; and an outlet passageway extending 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. It is also 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.

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. In one form of a variable geometry turbine, an axially moveable wall member, generally referred to as a“nozzle ring”, defines one wall of the inlet passageway. The position of the nozzle ring relative to a facing wall of the inlet passageway is adjustable to control the axial width of the inlet passageway. Thus, for example, as gas flow through the turbine decreases, the inlet passageway width may be decreased to maintain gas velocity and optimise turbine output. In different arrangements, the nozzle ring is fixed and the shroud is axially moveable so as to control the axial width of the inlet passageway.

The nozzle ring is provided with a plurality of vanes which extend into the inlet and through respective apertures (“slots”) provided in a“shroud” defining the facing wall of the inlet passageway to accommodate movement of the nozzle ring with respect the shroud. Movement of the nozzle ring with respect the shroud also controls the degree to which the vanes project through the respective slots.

In known variable geometry turbochargers which employ vanes projecting through slots in a shroud a clearance is provided between the vanes and the edges of the slots to permit the vanes to travel through the slots even when thermal expansion of the vanes occurs as the turbocharger operates at relatively high temperatures (e.g. between about 300°C and about 900°C). Such clearances may, in certain circumstances, provide undesired gas flow paths where gas may flow between the vanes and the edges of the slots behind the shroud. Such undesired gas flow paths may reduce the efficiency of a turbocharger.

It is one object of the present invention to obviate or mitigate one or more disadvantages of known turbines, whether mentioned above or otherwise. It is also an object of the present invention to provide for an improved or alternative variable geometry turbine.

Summary of Invention

According to an aspect of the invention there is provided a variable geometry turbine comprising a turbine wheel supported for rotation about a turbine axis; an annular inlet passageway surrounding the turbine wheel and defined between a first radial inlet surface of a first wall member and a second radial inlet surface of a second wall member, at least one of said first and second wall members being moveable along the turbine axis to vary the size of the inlet passageway; an array of vanes extending across the inlet passageway, said vanes being connected to said first wall member; a plurality of cavities located in the second wall member or to a rear side of the second wall member, wherein the rear side of the second wall member is located axially beyond the second wall member with respect to the first wall member; an array of vane slots in the second radial inlet surface, complementary to said array of vanes, said vane slots being configured to receive said vanes and permit the vanes to travel into said plurality of cavities, to accommodate relative movement between the first and second wall members; wherein the plurality of cavities comprises a first cavity, and a second cavity separated from the first cavity by a separating portion.

The separating portion may be configured to impede the flow of fluid between the first cavity and the second cavity.

Said impeding the flow of fluid between the first cavity and the second cavity may be impeding the flow of a fluid between the first cavity and the second cavity at a location axially beyond the second radial inlet surface with respect to the first wall member.

The separating portion may be configured to inhibit direct flow communication between the first cavity and the second cavity, that is to say flow communication between the first cavity and the second cavity that does not occur via the inlet passageway.

The inhibited direct flow communication between the first cavity and the second cavity may be such that there is no direct flow path between the first cavity and second cavity in a circumferential direction around the turbine axis.

In some embodiments, the separating portion is such that the only fluid flow path between the first cavity and the second cavity is via the inlet passageway. The plurality of cavities are located to said rear side of the second wall member and wherein each cavity is defined by a third wall member.

The third wall member may form part of a turbine housing, which houses the turbine wheel.

The plurality of cavities may be circumferentially spaced around the turbine axis.

Each of the plurality of cavities may define a segment of an annulus around the turbine axis.

The plurality of cavities may comprise an array of cavities which correspond to the array of vanes. The number of cavities may be equal to the number of vanes, such that each vane may be received in a corresponding cavity.

There may be the same number of separating portions as there are cavities.

In some embodiments each cavity may correspond to more than one vane, such that, in use, each cavity receives more than one vane.

In other embodiments there may be 2, 3, 4, 5 or 6 similar cavities. The cavities may each subtend an angle of approximately 180°, 120°, 90°, 72° or 60°. Of course, in other embodiments there may be any number of cavities, with any appropriate number of separating portions. The cavities may not all be similar, such that one or more cavities subtend a different angle to another cavity.

The shape of one or each of the cavities and the shape of one or each of the vanes may be complementary. The cross-sectional profile, in a plane perpendicular to the turbine axis, of the one or each relevant cavity may be complimentary to the cross-sectional profile, in a plane perpendicular to the turbine axis, of the one or each vane received by the one or each cavity.

That is, the cross-sectional profile or shape of a cavity may correspond to the cross- sectional profile or shape of a corresponding vane. The cross-sectional shape of the cavity and the cross-sectional shape of the vane may be substantially similar. The cross- sectional shape of the cavity may be concentric to the cross-sectional shape of the vane. The cross-sectional shape of the cavity may be a scaled version of the cross-sectional shape of the vane. The cross-sectional shape of a cavity may be said to be substantially geometrically similar to the cross-sectional shape of its received vane.

Alternatively, or in addition, the axial profile of the one or each relevant cavity may be complimentary to the axial profile of the one or each relevant vane.

The second and/or third wall member may be formed by additive manufacturing, metal injection moulding, a sintering process, machining from solid, electrical discharge machining or investment casting.

The second wall member may comprise a shroud plate and the third wall member may comprise a cavity plate. Said shroud plate and cavity plate may be secured to one another.

The second wall member and third wall member may be unitarily formed. That is, they may be formed integrally as one piece.

The second wall member, third wall member and turbine housing may be unitarily formed. That is, the turbine housing may be formed such that it integrally comprises the second wall member and third wall member. That is, the second wall member, third wall member and turbine housing may formed as a single piece. One or each vane may have a first vane height measured, parallel to the turbine axis, from the first radial inlet surface to a distal end of the vane.

One or each cavity may have a first cavity depth measured, parallel to the turbine axis, from the second radial inlet surface to a distal end of the cavity.

The first cavity depth may be slightly greater than the first vane height. That is, the axial profile of the cavity may correspond to, but be slightly deeper than, the axial profile of a corresponding vane. In particular, the cavity shape may correspond to the shape of a single-height vane. The axial profile of the cavity may be substantially geometrically similar to the axial profile of its received vane.

One or each vane may have a first vane portion with the first vane height, and a second vane portion with a second vane height, the second vane height being smaller than the first vane height. Such a vane may be referred to as a flag-cut-out vane.

One or each cavity may have a first cavity portion configured to receive the first vane portion and having said first cavity depth, and a second cavity portion configured to receive the second vane portion.

The second cavity portion may have a second cavity depth measured, parallel to the turbine axis, from the second radial inlet surface to a distal end of the second cavity portion.

The second cavity depth may be i) smaller than the first cavity depth and ii) slightly greater than the second vane height. That is, the axial profile of the cavity may correspond to, but be slightly deeper than, the axial profile of a corresponding vane. In particular, the cavity shape may correspond to the shape of a multi-height (flag-cut-out) vane.

When the first and second wall members are in a fully closed configuration, there may be a clearance defined as the shortest distance between a vane and a wall of a corresponding cavity.

The clearance may be in the range 0.1 to 3% of a diameter of a leading edge of the vane. The nozzle vane leading edge diameter is measured in a plane perpendicular to the turbine axis and is twice the distance between i) a point on the leading edge of the vane which intersects the chord line (or camber line) of the vane, and ii) the turbine axis.

Optionally, the clearance between the vane and the wall of the corresponding cavity at any point on the surface of the vane may be in the range 0.1 to 3% of a diameter of a leading edge of the vane.

When said first and second wall members are in said fully closed configuration, a slot clearance, defined as the shortest distance between a vane and a wall of a corresponding cavity, at a location proximal to a vane slot of said corresponding cavity, may be in the range 0.1 to 0.3 % of the diameter of the leading edge of the vane.

A portion of the third wall member defining a first cavity may be substantially isolated from a portion of the third wall member defining a second cavity adjacent the first cavity. That is, the portion of the third wall member defining the first cavity is not directly connected to the portion of the third wall member defining the second cavity. As such, any distortion on the portion of the third wall member defining the first cavity member may not cause a distortion on the portion of the third wall member defining the second cavity. The portions of the third wall member may be connected via a portion of the second wall member.

An axial profile, in a direction parallel to the turbine axis, of one or each of the cavities may be complimentary to an axial profile, in a direction parallel to the turbine axis, of a received vane.

That is, a depth profile of a cavity may correspond to a height profile of a corresponding vane. The axial profile of the cavity and the axial profile of the vane may be substantially geometrically similar. The axial profile of the cavity may be concentric to the axial profile of the vane. The axial profile of the cavity may be a scaled version of the axial profile of the vane.

The variable geometry turbine may further comprise one or more anti-leakage surface features located on an inner wall of a cavity. Optionally, said one or more anti-leakage surface features may comprise a concavity. The anti-leakage surface features may comprise a plurality of concavities. According to another aspect of the invention there is provided a turbocharger comprising a variable geometry turbine according to any preceding claim. Alternatively, the turbine may form part of any appropriate turbomachine, such as a power turbine.

Brief Description of the Drawings

Specific embodiments of the present invention will now be described, by way of example only, with reference to the accompanying drawings in which:

Figures 1 and 2 depict a cross-sectional view of a turbocharger assembly;

Figure 3 depicts a nozzle and shroud of a turbocharger assembly;

Figures 4A and 4B schematically depict gas flow pathways in a conventional turbocharger assembly compared to an embodiment of the invention;

Figures 4C-E schematically depict, in a plane perpendicular to the turbine axis, three different configurations of a plurality of shroud cavities and associated separating portions;

Figure 5A-C depicts a cavity wall assembly according to an example implementation of the invention;

Figures 6A and 6B show cross-sectional and perspective views of a modified shroud plate which may form art of a turbine according to the present invention;

Figures 7A and 7B show cross-sectional and perspective views of another modified shroud plate which may form part of a turbine according to the present invention; Figures 8A and 8B show cross-sectional and perspective views of another modified shroud plate which may form part of a turbine according to the present invention

Figure 9 shows a perspective view of part of a further modified shroud ring in section;

Figure 10 shows a view of a portion of a turbine housing which may form part of a turbine in accordance with the present invention; and

Figures 11A and 11 B show a cross-sectional view of a cavity of two modified shroud plates which may form part of a turbine according to the present invention.

Detailed Description

Figures 1 and 2 illustrate a variable geometry turbocharger in accordance with an embodiment of the present invention. The variable geometry turbine comprises a turbine housing 1 and a compressor housing 2 interconnected by a central bearing housing 3. A turbocharger shaft 4 extends from the turbine housing 1 to the compressor housing 2 through the bearing housing 3. A turbine wheel 5 is mounted on one end of the shaft 4 for rotation within the turbine housing 1 , and a compressor wheel 6 is mounted on the other end of the shaft 4 for rotation within the compressor housing 2. The shaft 4 rotates about turbocharger axis 4a on bearing assemblies located in the bearing housing 3.

The turbine housing 1 defines an inlet volute 7 to which gas from an internal combustion engine (not shown) is delivered. The exhaust gas flows from the inlet volute 7 to an axial outlet passageway 8 via an annular inlet passageway 9 and the turbine wheel 5. The inlet passageway 9 is defined on one side by a face 10 of a radial wall of a movable annular wall member 1 1 , which may be referred to as a“nozzle ring” or“nozzle”, and on the opposite side by a second wall member 12 comprising an annual shroud which forms the wall of the inlet passageway 9 facing the nozzle ring 11. The shroud 12 covers a generally annular recess 12a in turbine housing 1. The recess 12a is not circumferentially continuous around the turbine axis. Instead, the recess comprises a plurality of cavities 13 separated by one or more separating portions (not shown). The shroud 12 comprises an annular shroud plate 12b (see Figure 3). The nozzle ring 1 1 comprises an array of circumferentially, equally (angularly) spaced inlet vanes 14 each of which extends across the inlet passageway 9. The vanes are orientated to deflect gas flow through the inlet passageway 9 towards the direction of rotation of the turbine wheel 5. Vanes 14 project through suitably configured vane slots 25 (see Figure 3) in the shroud plate of the shroud 12, and into the plurality of cavities 13, to accommodate movement of the nozzle 1 1.

The position of the nozzle ring 1 1 is controlled by an actuator assembly. An actuator (not shown) is operable to adjust the position of the nozzle ring 11 via an actuator output shaft (not shown), which is linked to a yoke 15. The yoke 15 in turn engages axially extending actuating rods 16 that support the nozzle ring 11. Accordingly, by appropriate control of the actuator (which may for instance be pneumatic or electric), the axial position of the actuating rods 16 and thus the nozzle ring 11 can be controlled. The speed of the turbine wheel 5 is dependent on the velocity of the gas passing through the annular inlet passageway 9. For a fixed rate of mass of gas flowing into the inlet passageway 9, the gas velocity is a function of the width of the inlet passageway 9, the width being adjustable by controlling the axial position of nozzle ring 11. Figure 1 shows the annular inlet passageway 9 fully open. The inlet passageway 9 may be closed to a minimum by moving the nozzle ring 11 (and hence the face 10 of the nozzle ring 1 1) towards the shroud 12. Closing the inlet passageway 9 causes the vanes 14 to be inserted through the vane slots in the shroud 12, and subsequently into the plurality of cavities 13.

The nozzle ring 11 has axially extending radial inner and outer flanges 17, 18 that extend into an annular flange cavity 19 provided in turbine housing 1. Inner and outer sealing rings 20, 21 are provided to seal the nozzle with respect to inner and outer surface of the flange cavity 19 respectively, whilst allowing the nozzle ring 1 1 to slide within the flange cavity 19. The inner sealing ring 20 is supported within an annual groove formed in the radially inner annular surface of the flange cavity 19 and bears against the inner annular flange 17 of the nozzle ring 11. The outer sealing ring 21 is supported within an annular groove formed in the radially outer annular surface of the flange cavity 19 and bears against the outer annular flange 18 of the nozzle ring 1 1. Gas flowing from the inlet volute 7 and to the outlet passageway 8 passes over the turbine wheel 5 and, as a result, torque is applied to the shaft 4 to drive the compressor wheel 6. Rotation of the compressor wheel 6 within the compressor housing 2 pressurises ambient air present in an air inlet 22 and delivers the pressurised air to an air outlet volute 23 from which it is fed to an internal combustion engine (not shown).

To help characterise the invention, directional language may be defined. In particular, the air inlet 22 may be considered the “front” of the turbocharger and the outlet passageway 8 may be considered the“back” or“rear” of the turbocharger. In this way, the location of parts may be described in terms of their relative position along the turbine axis 4a. For example, the shroud 12 may be described as“behind” or“to the rear of’ the nozzle ring 1 1. “Behind” in this instance should be taken to mean closer to the back of the turbocharger. By way of further example, the plurality of cavities 13 are disposed behind the shroud 12.

Similarly, the movement of parts may be described in terms of their relative movement along the turbine axis 4a. Movement towards the back of the turbocharger may be considered“backwards” movement and movement towards the front of the turbocharger may be considered“forwards” movement. For example, movement of the nozzle ring 11 towards the shroud 12 to reduce the volume of the inlet passageway 9 may be described as the nozzle ring 11 moving in a backwards direction.

In the presently described embodiment the nozzle ring 1 1 is located in front of the shroud, such that the nozzle ring is located on the bearing housing side of the shroud. In other embodiments the nozzle ring may be located behind the shroud such that the nozzle ring is on the turbine housing side of the shroud. Furthermore, in the present embodiment, the shroud is stationary and the nozzle ring is actuated. In other embodiments the nozzle ring may be stationary and the shroud may be actuated. Alternatively, both the nozzle ring and shroud may be actuated.

Figure 3 shows an example shroud 12 and nozzle ring 11 shown in isolation from the turbocharger housing. As already discussed, the shroud 12 comprises an annular plate (also referred to as a shroud plate) 12b comprising a radially extending shroud wall 24 provided with vane slots 25 for the receipt of vanes 14 of the nozzle ring 1 1.

In this example, the inlet passageway 9 is defined between the nozzle ring 1 1 and a face (or radial inlet surface) of the shroud 12 proximal to the nozzle ring. In Figure 3 the separating portion defining different cavities is not shown, but the plurality of cavities 13 occupy a volume extending axially from the opposite face 12c of the shroud 12, in a direction away from the nozzle 11 and inlet passageway 9. That is, the plurality of cavities 13 are disposed behind the shroud 12.

The radially inner periphery of the annular shroud wall 24 is formed with an axially extending flange 26, which extends in an inboard direction away from the inlet passageway 9 when the shroud 12 is in position in the turbine housing, and provides means for seating the inner periphery of the shroud 12 in the mouth of the plurality of shroud cavities 13. The radially outer periphery of the shroud plate 24 is formed with a grooved flange 27. The flange 27 extends axially inboard (i.e. behind) from the shroud plate wall 24 to a greater extent than the inner shroud flange 26, and defines an annular groove 28 around the radially outer periphery of the shroud. A snap ring 28a or the like may be located in groove 28 and a corresponding groove 28b in the turbine housing to locate the shroud plate within the turbine housing. These outer and inner flanges 26, 27 and snap ring 28a form an imperfect seal between the inlet passageway 9 and the plurality of shroud cavities 13 behind the shroud 12 such that gas flow from the inlet passageway to the plurality of shroud cavities via the radially inner and outer edges of the shroud 12 is reduced.

In use, the vanes 14 may change in size, for example due to thermal expansion as the turbine is exposed to exhaust gas at elevated temperature. Known turbochargers may be designed such that the vanes 14 and slots 25 of the shroud are sized and shaped such that there is a clearance between the vanes 14 and the edges of the slots 25 to permit thermal expansion of the vanes relative to the slots without the vanes fouling on the slots. Should such fouling occur, axial movement of the nozzle ring relative to the shroud may be inhibited or prevented, resulting in reduced effectiveness of the turbine. By way of example in relation to the clearance between the vanes 14 and the edges of the slots 25, as viewed in the axial direction (i.e. such that what is observed is a cross- section in the plane perpendicular to the axis), the vanes 14 and slots 25 may have substantially the same shape, but the vanes 14 will be smaller than the slots 25. As such, in most situations (for example at temperatures equal to and below a maximum operating temperature), there is a clearance gap between the vanes 14 and the slots 25. The applicant has found that the clearance gap acts as a possible flow pathway for gas within the inlet passageway 9. A gas flow pathway exiting the inlet passageway 9 via a clearance gap may also be known as a“leakage pathway” or simply“leakage”.

Leakage represents an undesirable gas flow pathway. A major factor in the efficiency of a turbomachine is the vector at which gas strikes the blades of the turbine wheel 5, which is in turn dependent on the gas flow velocity vector of gas in the inlet passageway 9. Undesirable gas flow pathways from the inlet passageway 9, such as leakage through clearance gaps can alter the gas flow velocity vector in the inlet passageway 9. For example, leakage can occur at the clearance gap at a particular location between one vane and slot, such that the leaked gas from the inlet passageway goes behind the shroud and then flows back into the inlet passageway at another location (either at the same vane/slot as the leakage originated, or at a different vane/slot). When the leaked gas flows back into the inlet passageway it will do so in a generally random direction. Because this random direction will be different to the velocity vector of gas in the inlet passageway, it will cause the resultant velocity vector of gas in the inlet passageway (i.e. velocity vector of gas in absence of leaked gas re-entering the inlet passageway plus velocity vector of gas re-entering the inlet passageway) to change (e.g. away from the gas flow velocity vector which is optimised for the turbine concerned), thereby resulting in a reduction in turbine efficiency. As such, undesirable gas flow pathways may cause significant losses in turbine efficiency.

The applicant has realised that by reducing pathways for leakage associated with the nozzle and shroud, the efficiency of a variable geometry turbine may be increased. This is achieved by the provision of a separating portion as described below.

Figures 4A and 4B schematically show a circumferential cross section (i.e. cross-section in a cylindrical plane which runs along a circumference about the turbine axis, and which runs parallel to the turbine axis) of a sector of the nozzle ring 1 1 and shroud 12, comparing that of a prior art turbine and that of an embodiment of the present invention. Axes are shown to clarify the orientation of the nozzle ring 1 1 and shroud 12 with regard the turbine axis 4a and circumferential direction C. In both examples, an inlet passageway 9 is defined as above and the nozzle ring 1 1 and/or the shroud 12 may be moved with respect each other to alter the size of the inlet passageway 9. The shroud 12 comprises an array of vane slots 25 (two of which are shown) configured to receive a corresponding array of vanes 14a, 14b (again, two of which are shown) which form part of the nozzle ring 11. Gas may flow through these vane slots, as indicated by arrows 45, for example, through clearance gaps between vane 14a and the edge of a vane slot 25 as described above.

In the prior art turbocharger, behind the shroud there is a shroud cavity 130 into which the vanes may be received having passed through the vane slots. The shroud cavity is conventionally annular, extending circumferentially around the turbine axis 4a and having a cross-section perpendicular to the circumference which is substantially uniform - put another way, no feature exists in the shroud cavity 130 which modifies or impedes circumferential flow of gas within the shroud cavity.

After gas has flowed (or leaked) from the inlet passageway 9 to the cavity 130, it may continue to travel through the cavity 130. For example, it may travel in a circumferential direction C. It may travel from a first vane 14a to a second vane 14b via the cavity. Gas travelling between vanes 14a, 14b via the cavity represents a flow pathway for leakage, shown by an arrow 46. This may also be known as“vane to vane leakage”, or“cross talk”, indicating that vanes are in direct communication with each-other via the cavity. Vane to vane leakage may result in gas moving from a low-pressure region to a high- pressure region, for example from a high pressure side of a first vane 14a to a low pressure side of a second vane 14b. Vane to vane leakage is a sub-optimal flow pathway and, as discussed above, may contribute to a loss of efficiency in a variable geometry turbine.

It will be appreciated that whilst the example discussed above relates to a flow path for leaked gas between adjacent vanes, such a flow path for leaked gas may be between one vane and any other vane, after completing part of a circuit of the circumferential cavity 130 or more than one circuit of the cavity. The flow path may even be between one vane and the same vane after completing one or more circuits of the circumferential cavity 130.

Figure 4B shows how vane to vane leakage may be inhibited or prevented, in accordance with the present invention, through the provision of a separating portion 48. The separating portion 48 is provided behind the shroud 12, such that, instead of a single cavity 130, a plurality of cavities 13 are defined. For example, figure 4B depicts a single separating portion 48 defining two cavities 41 , 42.

The separating portion 48 acts to impede the flow of gas behind the shroud 12. In this example, the direction of impeded gas flow is in a circumferential direction C around the turbine axis 4a, for example in a direction from a first vane 14a to a second vane 14b. As such, the flow of gas between a first cavity 41 (which, in use, may contain the first vane) and a second cavity 42 (which, in use, may contain the second vane) may be impeded. In this instance, any gas flow entering the first cavity 41 may travel within first cavity 41 , but may not travel to the second cavity 42 without first travelling through the inlet passageway 9. In other words, the separating portion may remove or restrict direct communication between the first cavity 41 and the second cavity 42. If the first cavity 41 holds the first vane 14a and the second cavity 42 houses the second vane 14b, this also removes or restricts direct communication between the first and second vanes 14a, 14b. Direct communication should be understood as travel from a first volume to a second volume without travelling through any intermediate volume. In this way, reduced turbine efficiency due to cross-talk leakage between the vanes is minimised or eliminated.

A single separating portion 48 may be used to impede gas flow between a first and second vane 14a, 14b. The applicant has found that, by having at least two (i.e. a plurality) of cavities 13 separated by at least one separating portion 48, the efficiency of a variable geometry turbine may be increased. This is attributed at least in part to reduced cross-talk between vanes 14a, 14b.

In the above examples, the separating portion may be said to impede the flow of fluid between the first cavity and the second cavity. Said is impeding the flow of a fluid between the first cavity and the second cavity may be said to occur at a location axially beyond the second radial inlet surface with respect to the first wall member. In other words, the impeding of the flow of fluid between the first cavity and the second cavity is an impeding of flow which may happen to the rear of the inlet surface of the second wall member (in this case, the front face of the shroud 12)

The applicant has also found that, in some instances, it may be beneficial to provide a specific number of cavities and/or separating portions. In one example, shown in Figure 4C, two separating portions 48a may be used to provide two cavities 42a. These two cavities 42a may each subtend, for example, 180° such that two equal semi-circular cavities are provided. In another example, shown in Figure 4D, four separating portions 48a may be used to provide four cavities 42a. These four cavities 42a may each subtend, for example, 90° such that four quadrant cavities are provided. The number of cavities and separating portions may be chosen depending on the user’s needs, for example to optimise the pressure distribution around different vanes. It will be appreciated that, in the examples shown in Figures 4c and 4d, multiple vanes may be received in each of the cavities.

Figure 4e shows an embodiment including 16 cavities 42a and associated separating portions 48a. In this embodiment, the nozzle ring of the turbine may include 16 vanes such that each vane (not shown) is received in its own cavity. That is to say, each cavity 42a, in use, may receive only one vane.

It should be understood that, depending on the specific operating conditions required and/or the applicability of different manufacturing methods, the separating portion 48 may be provided in a number of ways. The separating portion 48 may be associated with the shroud 12 - e.g. it may form part of the shroud or be mounted to the shroud. Alternatively or additionally, the separating portion 48 may be provided by a portion of the turbine housing 1. Alternatively or additionally, the separating portion 48 may be associated with an additional body, for example, a cavity wall which may be provided behind the shroud 12 as described in more detail below. In each case, the separating portion 48 and the shroud 12 and/or cavity wall and/or housing 1 may be a single piece or may be separate pieces joined together. In some cases, it is desirable for a separating portion 48 to extend completely between the cavities 41 , 42 it separates, for example to completely remove direct communication between the two cavities. However, in other situations it may be beneficial for the separating portion 48 to extend only partially between the cavities 41 , 42 it separates, so as to allow reduced or inhibited communication between the two cavities 41 , 42, for example if a small amount of communication is required for pressure equalizing.

As mentioned above, a cavity wall may be provided behind the shroud. A cavity wall may be used to further define the cavities 13, for example the axial extent of the cavities 13 in a backwards direction from the shroud 12 and/or the radial extent of the cavities 13 e.g. between the inner and outer peripheries of the shroud 12. However, the cavity wall is not an essential feature. In some cases (for example for reduced cost and/or ease of manufacture) it is beneficial to omit a cavity wall and instead provide a separating portion 48 using, for example, features in the turbine housing 1.

Alternatively, a cavity wall may be provided separate to the turbine housing 1. In this case, the cavity wall may be of unitary construction with the shroud 12 and/or separating portion 48. Alternatively, the cavity wall may be a separate piece that can be joined to the shroud 12 and/or separating portion 48 and/or housing 1 through any suitable method as known in the art, for example stamping, press fitting, etc. A benefit of providing a separate cavity wall is that it allows increased flexibility in the manufacturing methods used, for example sintering, injection moulding, stamping assembly, machining from solid, electrical discharge machining, investment casting etc. Post machining may also be performed as required, for example to further define critical features.

Another benefit of providing a separate cavity wall is that it allows increased flexibility in the shape and/or orientation of cavities 13 and separating portions 48. As described above, different numbers of cavities and separating portions may be chosen depending on the user’s needs. Different cavity shapes may also be provided, aided by the provision of cavity wall which may further define the size and shape of the cavities. A particular example will be described with reference to Figures 5A-C. Figures 5A-C show a nozzle 1 1 and cavity wall assembly 50 for use in a variable geometry turbine of a turbocharger. The turbine comprises a turbine wheel 5 in a turbine housing 1. In this example, there is provided a cavity wall assembly 50 comprising a one-piece cavity wall 52 and separating portions 48 which cooperate to form an array of cavities 13. The number of cavities 13 is equal to the number of vanes 14 of the nozzle ring. The cavity wall assembly 50 in this instance is mounted behind the shroud 12 such that each vane 14 is received by a corresponding cavity 13 as the nozzle ring moves towards the shroud. The cavity wall assembly 50 may be considered to constitute a third wall member within the language of the claimed invention.

In Figures 5A-C, the cavities 13 are shaped such that they have substantially the same shape as the vanes 14. The cavities 13 also have substantially the same cross-sectional shape (in a plane perpendicular to the axis) as the vane slots 25 in the shroud 12. The cavities 13 are larger in volume than the vanes 14 so as to allow clearance between each vane 14 and its respective cavity 13.

A benefit of the example shown in Figures 5A-C is that, by shaping the cavities 13 such they correspond to that of their received vane, the volume of each cavity 13 is as low as possible. With a reduced volume of cavity, a smaller volume of gas may be retained in a cavity 13 at any given time. Given usual flow conditions, this results a reduced rate of leakage gas flow to/from the cavity 13 from/to the inlet passageway 9. As a result, shaping the cavities 13 such that they correspond to that of their received vane, to reduce cavity volume, can beneficially increase the efficiency of a variable geometry turbine.

An additional benefit of the above example is that, by selecting the shape of the cavities 13, the pressure distribution around each vane 14 may be optimised. The applicant has found that, in one example, it may be beneficial to provide a non-uniform cavity clearance between a vane 14 and a cavity 13. Cavity clearance is the shortest distance between the edge of a vane 14 and the edge of a cavity 13 for a particular point on the vane 14. It should be understood that, while a smaller cavity clearance may result in increased turbocharger efficiency, a larger cavity clearance may result in increased robustness. As such, cavity clearance may be optimised depending on the required usage conditions. The applicant has determined that in some applications, providing a maximum cavity depth (i.e. in the axial direction) that is slightly greater than the maximum vane depth (again, in the axial direction) may be beneficial. For example, the maximum cavity depth may preferentially be between 1.2 to 2 times the maximum vane depth. The maximum cavity depth may be measured relative to the inlet surface of second wall member. The maximum vane depth may be measured relative to the inlet surface of the first wall member.

In the example shown in Figures 5A-C, the separating portions 48 in the cavity wall assembly 50 are of complementary shape to that of the vanes 14 that the particular separating portion is located between. Alternatively, in the schematic shown in Figure 4B, the separating portion is depicted as a rectilinear object. However, it should be understood that the separating portion may take any shape. The separating portion may be further shaped to affect the gas flow. For example, the exterior of the separating portion may be curved to impart a change in direction in any leaked gas received by the cavity, and/or to reduce turbulent flow.

Figures 6 to 11 show several different embodiments of the invention in which the plurality of cavities and separating portions are formed in different ways.

Figures 6A and 6B show a modified shroud plate 54 which forms part of a turbine according to the present invention. The shroud plate 54 is mounted in the turbine housing in the same manner as discussed above in relation to Figures 2 and 3, utilising circumferential groove 28. The shroud plate 54 comprises vane slots 25 and, associated with each vane slot, a cavity wall member 56 which is integral with the shroud plate 54. Each cavity wall member 56 defines a cavity 13 behind its respective vane slot 25. An example of a manufacturing method which may be used to form the shroud plate is Metal Injection Moulding (MIM) or sintering process, with post machining, if required.

Figures 7A, 7B, 8A and 8B show modified shroud plates 74, 84 similar to the modified shroud plate 54 shown in Figures 6A and 6B, wherein the shape of the modified shroud plates 74, 84 are shaped so as to conform to the shape of the vanes. Each modified shroud plate 74, 84, comprises vane slots 25 and, associated with each vane slot, a cavity wall member 56 which is integral with the shroud plate 74, 84. Each cavity wall member 56 defines a cavity 13 behind its respective vane slot 25. The array of cavities 13 defined by the cavity wall members 56 accommodate movement of a nozzle 11 comprising an array of vanes 14. In Figures 7A, 7B and 8B, the nozzle 11 is shown fully inserted into the shroud plate 74, 84 such that the vanes 14 cannot extend further through the slots 25, and the axial width of the annular inlet passageway between opposing faces of the nozzle 11 and the shroud plate 74, 84 is a minimum. Such a position may be referred to as fully closed. For completeness, the view shown in Figure 8A is an exploded view.

In the example depicted in Figures 7 A and 7B, the vanes have a height profile which is substantially single-height. That is, the height 70 of the vanes is substantially constant across the vane, for example from first edge 78 of the vane 14 to a second edge 79 of the vane 14. The height profile may be referred to as an axial profile.

In the example depicted in Figures 8A and 8B, the vanes have a height profile which is multi-height. That is, the height (80F, 80S) of the vanes is not constant across the vane. In this example, a first vane portion 14F has a larger (or greater) height than a second vane portion 14S of the vane 14. The first vane portion 14F is closer to a first (leading) edge 78 of the vane and the second portion 14S is closer to a second (trailing) edge 79 of the vane. The vanes may be considered to be equivalent to a single-height vane with a portion of vane removed or‘cut-out’ (i.e. the second vane portion 14S is smaller in height due to the‘cut-out’). Such multi-height vanes may be referred to as flag-cut-out vanes, or vanes having a flag-cut geometry.

Figures 7A and 7B show a modified shroud plate 74 shaped so as to conform to the shape of vanes with a height profile which is substantially single-height. The vanes 14 have a vane height 70 measured, parallel to the turbine axis, from a face 10 of the nozzle 1 1 , to a distal end 71 of the vane 14. The cavities 13 have a cavity depth 72, measured parallel to the turbine axis, from a face of the shroud plate 74 which opposes the face 10 of the nozzle 11 , to a distal end 73 of the cavity 13. In this arrangement, the cavities have a cavity depth 72 which is substantially constant across the cavity 13 i.e. the cavities 13 have a corresponding single-depth height profile. The height profile of the cavities may also referred to as a depth profile or an axial profile. In fact, the cavity depth 72 may have some local changes in height, for example bevelling around the corners of the cavity 13, but generally has a cavity depth 72 which is substantially constant, particularly at the portion of the cavity which is axially above the camber line of the received vane.

Figures 8A and 8B show a modified shroud plate 84 shaped so as to conform to the shape of vanes with a height profile which is substantially multi-height (or stepped). The vanes 14 have a first vane portion 14F with a first vane height 80F and a second vane portion 14S with a second vane height 80S. The first and second vane heights 80F, 80S are measured, parallel to the turbine axis, from a face 10 of the nozzle 11 , to a distal end of the first vane portion 14F and second vane portion 14S, respectively.

In this arrangement, the cavities have a depth which is not constant, but changes from the first (leading) edge 78 to the second (trailing) edge 79 in correspondence with the height profile of the vanes 14. As such, the cavities 13 have a first cavity portion 13F arranged to receive the first vane portion 14F, and a second cavity portion 13S arranged to receive the second vane portion 14S. The first cavity portion 13F has a first cavity depth 82F, measured parallel to the turbine axis, from a face of the shroud plate 84 which opposes the face 10 of the nozzle 11 , to a distal end of the first cavity portion 13F. The second cavity portion 13S has a second cavity depth 82S, measured parallel to the turbine axis, from a face of the shroud plate 84 which opposes the face 10 of the nozzle 1 1 , to a distal end of the second cavity portion 13S.

In fact, it can be seen in Figures 8A and 8B that the second vane portion 14S of the gradually increases from the second edge 79 towards the first vane portion 14F. That is, the second vane portion 14S is lower in height than the first vane portion 14F, but has some local changes in height. The second cavity portion 13S is correspondingly shaped to have a similar (or corresponding) height profile to that of the second vane portion 14S. Furthermore, the first cavity portion 13F has a depth generally equal to the first cavity depth 82F, but with some local changes in depth. Such local changes in depth may be due to local changes in height of the first portion 14F of the vane 14, may be due to machining tolerances, and/or may be selected to control gas flow within the cavity (e.g. with curved corners rather than angular). It should be understood that in other embodiments any height profile of vane may be used, and the shape of the cavities in the shroud plate may be chosen to generally correspond to said height profile. The depth profile of the cavities may be chosen to be similar to, but slightly greater than, the height profile of vanes such that, upon being received by the cavities, the vanes do not touch the cavity walls. For example, the maximum depth of the cavities may be 1.0 to 1.2 times the maximum height of the vanes. The height profile of the vanes may be measured along the camber line of the vane and the depth profile of the cavity may be measured along the portion of the cavity which is axially above the camber line of the vane.

The height/depth profile of cavities 13 and vanes 14 may be considered in a plane parallel to the turbine axis. That is, a height or depth profile may be considered as the change in the axial extent of a vane or cavity. Alternatively, the height/depth profile of cavities 13 and vanes 14 may be considered in a surface which is parallel to the turbine axis and which is curved to follow the camber line of the vane.

In addition to the depth profile of the cavities 13 being made conformal to the height profile of the vanes 14, it may also be beneficial to select the cross-sectional shape/area of the cavities such that they correspond to the cross-sectional shape/area of the vanes, as described further above. Such a cross-sectional shape/area may be measured in a plane perpendicular to the turbine axis (i.e. comprising the radial and circumferential extent of the turbocharger).

An example of a manufacturing method which may be used to form the shroud plates illustrated in Figures 7A, 7B, 8A and 8B is injection moulding, for example Metal Injection Moulding (MIM), a sintering process, additive manufacturing, or investment casting, with post machining, if required.

In the above examples described with reference to Figures 7A, 7B, 8A and 8B, the shape of the cavity is generally matched to the shape of the vane (both in terms of axial height and cross-sectional shape/area). In this way, the excess volume of the cavity (i.e. the volume of a cavity not filled by a vane when in the fully closed position) is minimized. Beneficially, by reducing this excess volume, leakage may be reduced (for example, by minimising the volume, any pressure differential either side of the vane slots may be equalised more quickly therefore reducing leakage through the vane slots into the cavity). By selecting the cavity shape to generally match the vane shape, the pressure distribution around each vane 14 may be optimised, which may additionally reduce leakage.

The amount of excess volume may be quantified considering clearances within the cavity. Clearance is the shortest distance between the edge of a vane and the edge of a cavity at a given point on the vane. In a shroud plate with a cavity shape generally matching the vane shape, the clearance between a vane and a cavity is typically substantially constant when the vane is fully received by the cavity.

Figure 7B illustrates a height clearance 75 between the distal end 71 of the vane 14 and the distal end 73 of the cavity. Figure 7B also illustrates a radial clearance 76 between an edge of the vane 14 and a side wall of the cavity 13. The radial clearance 77 and height clearance 75 are located in the main body of the cavity and may be referred to as cavity clearances. Figure 7B also illustrates a slot clearance 77 between an edge of the vane proximal to the face 10 of the nozzle 11 and a slot 25.

By minimizing clearances 75, 76, 77, the excess volume may be further reduced, thereby reducing leakage. Small slot clearances 77 may be particularly desirable so as to reduce the area between the vane and the walls of the slot/cavity through which gas may travel between the cavity and the annular inlet. However, non-zero clearances 75, 76, 77 are typically desired in order to allow for part movement and/or thermal expansion. Additionally, adequate clearance may also be beneficial to allow for machining tolerances. As is shown in the example in Figure 7B, the slot clearance may be smaller than the cavity clearances (height clearance and/or radial clearance). However, in other embodiments, this need not be the case.

The applicant has found that cavity clearances 75, 76 of between 0.2 - 2 % (and, particularly, 0.3 - 1.5%) of the nozzle vane leading edge diameter may be desirable. The nozzle vane leading edge diameter is measured in a plane perpendicular to the turbine axis and is twice the distance between i) a point on the leading edge of the vane which intersects the chord line (or camber line) of the vane, and ii) the turbine axis. For example, in combination with a known nozzle, a cavity clearance in the range 0.3 to 1.5 mm may be desirable. A larger cavity clearance may be desirable for larger vanes, having a larger leading edge diameter. Such a clearance may allow for sufficient thermal expansion and/or part movement in use, while also sufficiently optimising the pressure distribution around each vane and/or reducing leakage.

Cavity clearances 75, 76 may be substantially constant around the vane 14. Alternatively, the cavities 13 may be shaped such that the cavity clearances 75, 76 are different in different portions of the cavity 13. For example, the radial clearance 76 may be different to the height clearance 75, and/or the clearance 75, 76 in the first portion 13F of the cavity may be different to the clearance 75, 76 in the second portion of the cavity 13S.

The applicant has found that slot clearances in the range 0.1 - 0.3 % (and, preferably, 0.1 - 0.2%) of the nozzle vane leading diameter may be desirable. For example, in combination with a known nozzle, a slot clearance in the range 0.1 to 0.2 mm may be desirable. A larger slot clearance may be desirable for larger vanes, having a larger leading edge diameter. The slot clearance 77 may be substantially constant around the slot 25. Alternatively, the slot clearance 77 may be different in different positions around the slot. For example, a larger slot clearance 77 may be provided around the leading edge compared to the trailing edge of the vane 14.

While clearances have been described with reference to Figure 7 A and 7B, any of the shroud plates described herein have clearances selected as described above. For example, the shroud plate 84 shown in Figure 8A and 8B may have a first cavity clearance in the first cavity portion 13F and a second cavity clearance in the second cavity portion 13S. The first and second cavity clearance may be selected to be in the ranges described above. The first and second cavity clearances may be substantially equal, or may be selected such that the first cavity clearance is different to the second cavity clearance, or even such that the clearance changes throughout the cavity.

In the above examples, the face of the shroud plates 74, 84 which includes the vane slots 25 may be referred to as a second wall member. The cavity wall members 56 may be referred to as a third wall member. Figures 7A to 8B show modified shroud plates 74, 84 which are unitarily formed. That is, the cavity wall members 56 are integral to the shroud plates 74, 84.

Figure 9 shows a modified shroud plate 94 which is similar to that shown in Figures 6A to 7B. The difference between the shroud plate 94 in this Figure and the shroud plates 54, 74 in Figures 6A and 6B and 7A and 7B is that the cavity wall members 60 which define each cavity 13 are formed as a separate piece (which may be referred to as a cavity plate 63) to the shroud plate portion 62 which includes the vane slots 25. The cavity plate 63 which includes the cavity wall members 60, and the shroud plate portion 62 may be joined or secured to one another in any appropriate manner. One such example is stamping. A modified shroud plate similar in shape to the shroud plate in Figures 8A and 8B may be formed in a similar manner, by securing a differently shaped cavity plate to the shroud plate portion.

In the examples described above and with respect to Figures 6A to 9, the cavity wall members 56, 60 are substantially isolated or separate from one another. That is, the cavity wall member 56, 60 of a first cavity 13 is not in direct contact with (or mechanically connected to) a cavity wall member 56, 60 of an adjacent cavity 13. Rather, they are connected indirectly via the annular face of the shroud plate 54, 74, 84, 94.

In use, a cavity wall member 56, 60 may experience tension and/or distortion, for example due to thermal expansion and/or vibrations. By providing cavity wall members 56, 58 which are substantially isolated, any tension/distortion experienced by a first cavity wall member 56, 58 will not affect an adjacent cavity wall member (or, in fact, any other cavity wall member). As such, even if a first wall member (and hence cavity 13) distorts in shape, adjacent/other cavity wall members (and hence adjacent/other cavities) should not be affected. In addition or alternatively, tension/distortion experienced by each of the cavity wall members will be substantially be prevented from having a cumulative effect on the overall tension/distortion experienced by the shroud plate. Advantageously, this may increase the durability of the turbocharger.

Figure 10 shows another embodiment of the invention. In this case, the cavities 13 for receiving each vane are formed by a cavity wall portion 64 of the turbine housing 1 a. That is, the cavity wall portion 64 which defines the cavities 13 is unitarily formed with the turbine housing 1a. As such, in this embodiment the cavities 13 are integral to the turbine housing 1a. Furthermore, the vane slots (formed in the previously referred to second wall member) are also unitarily formed with the turbine housing. Such an arrangement, in which the cavity wall member and/or vane slots are unitarily formed with the turbine housing, may further reduce leakage within the turbocharger, for example by reducing the number of seals between separate components. Additionally, by forming such features integrally/unitarily, this means that no relative rotation between the features is possible. This may help to prevent misalignment between such features. Such a turbine housing 1a may be formed in any appropriate manner, including additive manufacturing, MFS (machined from solid) and/or EDM (electrical discharge machining). The cavity wall portion 64 may be referred to as a third wall member which forms part of the turbine housing 1 a.

In some of the above-described embodiments the plurality of cavities are located in the second wall member. For example, in the embodiment shown in Figures 6A and 6B, the shroud plate 54 can be said to be the second wall member, in which cavities 13 are located. Likewise, in Figure 9, the two-piece shroud plate 58 can be said to be the second wall member, in which cavities 13 are located.

Alternatively, in some of the above-described embodiments the plurality of cavities are located to a rear side of the second wall member. For example, in Figures 5A-5C the plurality of cavities 13 are located in the cavity wall assembly 50. The cavity wall assembly 50 is located to the rear side (or rear) of the second wall member defined by the shroud 12. Likewise, in Figure 10 the plurality of cavities 13 are located in the cavity wall portion 64 of the turbine housing 1 a, which, although not shown in the figure, is located to the rear side (or rear) the shroud of the turbine.

In each of the embodiments discussed in the above paragraph, the rear side of the second wall member (shroud) is located axially beyond the second wall member with respect to the first wall member (nozzle ring). In other words, the rear side of the second wall member is located at a greater axial distance from the first wall member than the second wall member itself. In the embodiments discussed above the array of vane slots in the second radial inlet surface is complementary to the array of vanes of the nozzle ring. The vane slots are configured to receive the vanes and permit the vanes to travel into the plurality of cavities. In this way the vane slots accommodate relative movement between the first and second wall members (i.e. relative movement between the first and second wall members such that the axial spacing between the first and second wall members decreases), in use.

Additional features may be provided to the turbocharger to control gas flow. For example, anti-leakage surface features may be provided in the cavities to resist gas flow within the cavities. Such anti-leakage surface features may comprise a plurality of concavities configured to generate vortices in a fluid as described in PCT/GB2020/050416.

Figures 1 1A and 1 1 B depict a cross-sectional view of a cavity 13 of two example modified cavity wall members 110, 112. Both cavity wall members 110, 1 12 comprise anti-leakage surface features 114, 116 provided on an inner wall of the cavity 13. The anti-leakage surface features may act to inhibit flow of gas between a received vane and the cavity wall.

In the example in Figure 11 A, the cavity wall member 1 10 comprises anti-leakage surface features 1 14 in the form of recesses or concavities. In this particular example, although not necessarily the case in other embodiments, the anti-leakage features are directional. In particular, the anti-leakage surface features 114 comprise concavities which are angled, with respect the slot 25, to inhibit the flow of fluid from a distal end of the cavity towards the slot 25. Different directionalities may be employed in other example arrangements to inhibit the flow of fluid in other directions.

In the example in Figure 11 B, the cavity wall member 1 10 comprises anti-leakage surface features 1 16 which are generally hemispherical projections.

The anti-leakage surface features 114, 1 16 of each of the examples generally encourage the formation of vortices in a fluid so as to inhibit gas flow. In other applications, other shapes and geometries of anti-leakage features may be used. In some applications, anti-leakage surface features may additionally or alternatively be provided proximal to the slot 25.

Although the turbine discussed above forms part of a turbocharger, it will be appreciated that a turbine according to the present invention may form part of any appropriate turbo machine.

While gas has been specifically referred to in the above examples, it will be understood that the invention applies equally to a turbine which operates in conjunction with a liquid.

While a number of manufacturing methods have been named, for example, sintering, injection moulding, stamping assembly, MFS, EDM, investment casting, additive manufacturing, the components described above may be fabricated by a variety of methods as known in the art. Post machining may be performed as required, for example, to further define critical features.

The components described above may be fabricated from any appropriate material or combination of materials. A key consideration when determining which material(s) to use for fabrication will be whether the material can maintain its integrity in the operating environment of the turbine.

Additive manufacturing may make it possible to manufacture the multi-cavity shroud (e.g. including the second and/or third wall member) in a single piece. As discussed above and with reference to Figure 10, the multi-cavity shroud may also be manufactured with the turbine housing in a single piece. There are undercuts in the design that may make it challenging to cast without an expensive multi-piece casting tool. It may be possible to fabricate the multi-cavity shroud from many simple components, but the fabrication would introduce cost and dimensional variability that may be undesirable. Machining the part from billet may not be practical in some embodiments because of the thin, deep cavities. Beneficially, by manufacturing the shroud in a single piece, additional gas leakage pathways may be removed. Additionally/alternatively, clearances may be selected with more accuracy leading to enhanced resilience of the turbocharger.

Claims

1. A variable geometry turbine comprising:

a turbine wheel supported for rotation about a turbine axis; an annular inlet passageway surrounding the turbine wheel and defined between a first radial inlet surface of a first wall member and a second radial inlet surface of a second wall member, at least one of said first and second wall members being moveable along the turbine axis to vary the size of the inlet passageway;

an array of vanes extending across the inlet passageway, said vanes being connected to said first wall member;

a plurality of cavities located in the second wall member or to a rear side of the second wall member, wherein the rear side of the second wall member is located axially beyond the second wall member with respect to the first wall member;

an array of vane slots in the second radial inlet surface, complementary to said array of vanes, said vane slots being configured to receive said vanes and permit the vanes to travel into said plurality of cavities, to accommodate relative movement between the first and second wall members;

wherein the plurality of cavities comprises a first cavity, and a second cavity separated from the first cavity by a separating portion.

2. The variable geometry turbine of claim 1 , wherein the separating portion is configured to impede the flow of fluid between the first cavity and the second cavity.

3. The variable geometry turbine of claim 2, wherein said impeding the flow of fluid between the first cavity and the second cavity is impeding the flow of a fluid between the first cavity and the second cavity at a location axially beyond the second radial inlet surface with respect to the first wall member.

4. A variable geometry turbine according to any preceding claim wherein the plurality of cavities are located to said rear side of the second wall member and wherein each cavity is defined by a third wall member.

5. A variable geometry turbine according to claim 4, wherein the third wall member forms part of a turbine housing, which houses the turbine wheel.

6. A variable geometry turbine according to any preceding claim, wherein the plurality of cavities are circumferentially spaced around the turbine axis.

7. A variable geometry turbine according to any preceding claim, wherein the plurality of cavities comprises an array of cavities which correspond to the array of vanes and the number of cavities is equal to the number of vanes, such that each vane is received in a corresponding cavity.

8. The variable geometry turbine of any preceding claim, wherein the shape of one or each of the cavities and the shape of one or each of the vanes are complementary.

9. The variable geometry turbine of any preceding claim, wherein the cross- sectional profile, in a plane perpendicular to the turbine axis, of one or each of the cavities is complimentary to the cross-sectional profile, in a plane perpendicular to the turbine axis, of the one or each vane received by the one or each cavity.

10. The variable geometry turbine according to any preceding claim wherein the second and/or third wall member are formed by additive manufacturing, metal injection moulding, a sintering process, machining from solid, electrical discharge machining or investment casting.

1 1. The variable geometry turbine according to claim 4 or any claim depending on claim 4, wherein the second wall member comprises a shroud plate and the third wall member comprises a cavity plate, said shroud plate and cavity plate being secured to one another.

12. The variable geometry turbine of any of claims 4 to 10, wherein the second wall member and third wall member are unitarily formed.

13. The variable geometry turbine of claim 12, wherein the second wall member, third wall member and turbine housing are unitarily formed.

14. The variable geometry turbine of any preceding claim, wherein: one or each vane has a first vane height measured, parallel to the turbine axis, from the first radial inlet surface to a distal end of the vane;

one or each cavity has a first cavity depth measured, parallel to the turbine axis, from the second radial inlet surface to a distal end of the cavity; and

the first cavity depth is slightly greater than the first vane height.

15. The variable geometry turbine of claim 14, wherein:

one or each vane has a first vane portion with the first vane height, and a second vane portion with a second vane height, the second vane height being smaller than the first vane height;

one or each cavity has a first cavity portion configured to receive the first vane portion and having said first cavity depth, and a second cavity portion configured to receive the second vane portion;

the second cavity portion has a second cavity depth measured, parallel to the turbine axis, from the second radial inlet surface to a distal end of the second cavity portion; and

the second cavity depth is i) smaller than the first cavity depth; and ii) slightly greater than the second vane height.

16. The variable geometry turbine of any preceding claim, wherein, when said first and second wall members are in a fully closed configuration, there is a clearance defined as the shortest distance between a vane and a wall of a corresponding cavity, and wherein the clearance is in the range 0.1 to 3% of a diameter of a leading edge of the vane, and, optionally, wherein the clearance between the vane and the wall of the corresponding cavity at any point on the surface of the vane is in the range 0.1 to 3% of a diameter of a leading edge of the vane.

17. The variable geometry turbine of claim 16, wherein when said first and second wall members are in said fully closed configuration, a slot clearance, defined as the shortest distance between a vane and a wall of a corresponding cavity, at a location proximal to a vane slot of said corresponding cavity, is in the range 0.1 to 0.3 % of the diameter of the leading edge of the vane.

18. The variable geometry turbine of claim 4, or any claim dependent thereon, wherein a portion of the third wall member defining a first cavity is substantially isolated from a portion of the third wall member defining a second cavity adjacent the first cavity.

19. The variable geometry turbine of any preceding claim, wherein an axial profile, in a direction parallel to the turbine axis, of one or each of the cavities complimentary to an axial profile, in a direction parallel to the turbine axis, of a received vane.

20. The variable geometry turbine of any preceding claim further comprising one or more anti-leakage surface features located on an inner wall of a cavity, and, optionally, wherein said one or more anti-leakage surface features comprises a concavity.

21. A turbocharger comprising a variable geometry turbine according to any preceding claim.