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Title:
GAS TURBINE EXHAUST DIFFUSER HAVING FLOW GUIDING ELEMENTS
Document Type and Number:
WIPO Patent Application WO/2019/027661
Kind Code:
A1
Abstract:
A gas turbine (10) includes a turbine section (24), and a diffuser (30) located downstream of a row of turbine blades (26) of the turbine section (24). The diffuser (30) includes an annular duct (34) extending axially along a diffuser axis (32). The annular duct (34) is delimited radially by an outer wall (36) and an inner wall (38) which respectively define outer and inner boundaries of an exhaust flowpath. A number of diffuser struts (40) are circumferentially distributed within the annular duct (34). Each diffuser strut (40) extends from the outer wall (36) to the inner wall (38). A plurality of flow guiding elements (50) are circumferentially distributed within the annular duct (34). The flow guiding elements (50) are positioned on the inner wall (38) and extend radially therefrom into the exhaust flowpath. The flow guiding elements (50) have an axial location downstream of the row of turbine blades (26) and upstream of the diffuser struts (40). The number of flow guiding elements (50) is larger than the number of diffuser struts (40).

Inventors:
OROSA, John A. (11730 Cottonwood Circle, Palm Beach Gardens, Florida, 33410, US)
WALTKE, Ulrich (Wilhelmstr. 15, Mülheim an der Ruhr, Ruhr, DE)
Application Number:
US2018/042203
Publication Date:
February 07, 2019
Filing Date:
July 16, 2018
Export Citation:
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Assignee:
SIEMENS AKTIENGESELLSCHAFT (Werner-von-Siemens-Straße 1, München, München, DE)
SIEMENS ENERGY, INC. (4400 Alafaya Trail, Orlando, Florida, 32826-2399, US)
International Classes:
F01D25/30; F01D5/14; F01D9/04
Foreign References:
EP2672080A22013-12-11
GB2226600A1990-07-04
EP2378072A22011-10-19
US20060269399A12006-11-30
Other References:
None
Attorney, Agent or Firm:
BASU, Rana (Siemens Corporation- Intellectual Property Dept, 3501 Quadrangle Blvd. Ste. 230Orlando, Florida, 32817, US)
Download PDF:
Claims:
CLAIMS

1. A gas turbine (10) comprising:

a turbine section (24);

a diffuser (30) located downstream of a row of turbine blades (26) of the turbine section (24), the diffuser (30) comprising:

an annular duct (34) extending axially along a diffuser axis (32), the annular duct (34) being delimited radially by an outer wall (36) and an inner wall (38) which respectively define outer and inner boundaries of an exhaust flowpath,

a plurality of diffuser struts (40) circumferentially distributed within the annular duct (34), wherein each diffuser strut (40) extends from the outer wall (36) to the inner wall (38), and

a plurality of flow guiding elements (50) circumferentially distributed within the annular duct (34), the flow guiding elements (50) being positioned on the inner wall (38) and extending radially therefrom into the exhaust flowpath, the flow guiding elements (50) having an axial location downstream of the row of turbine blades (26) and upstream of the diffuser struts (40),

wherein the number of flow guiding elements (50) is larger than the number of diffuser struts (40).

2 The gas turbine (10) according to claim 1,

wherein the flow guiding elements (50) are axially located in a region in which flow is predicted to be already separated downstream of the turbine blades (26), during operation.

3. The gas turbine (10) according to claim 2,

wherein a local radial height (r) of one or more of the flow guiding elements (50) is greater than a predicted local radial height (s) of a flow separation region (70) over the inner wall (38), at a specified axial location.

4. The gas turbine (10) according to claim 3, wherein the local radial height (r) of said one or more of the flow guiding elements (50) is less than or equal to 20% above the predicted local radial height (s) of the flow separation region (70) over the inner wall (38), at the specified axial location.

5. The gas turbine (10) according to any of the preceding claims, wherein a local radial height (r) of one or more of the flow guiding elements

(50) lies in a range of 10 to 25% of a local radial distance (R) between the outer wall (36) and the inner wall (38), at a specified axial location.

6. The gas turbine (10) according to any of the preceding claims, wherein for one or more of the flow guiding elements (50), a local radial height at a leading edge (52) thereof is lesser than a local radial height at a trailing edge (54) thereof.

7. The gas turbine (10) according to any of the preceding claims, wherein the flow guiding elements (50) are axially located closer to a trailing edge position of the turbine blades (26) than to a leading edge position of the diffuser struts (40).

8. The gas turbine (10) according to any of the preceding claims, wherein a first axial distance (di) is defined between a leading edge position of the flow guiding elements (50) and a trailing edge position of the turbine blades (26), and a second axial distance (d2) is defined between the trailing edge position of the turbine blades (26) and a leading edge position of the diffuser struts (40),

the first axial distance (di) lying in the range of 5 to 25% of the second axial distance (d2).

9. The gas turbine (10) according to any of the preceding claims, wherein a first axial distance (di) is defined between a leading edge position of the flow guiding elements (50) and a trailing edge position of the turbine blades (26), the first axial distance (di) lying in the range of 5 to 25% of a hub side chord- length of the turbine blades (26).

10. The gas turbine (10) according to any of the preceding claims, wherein each flow guiding element (50) has a base (62) adjoining the inner wall (38) and a tip (64) opposite to the base (62),

wherein an axial length (h) of the base (62) is equal to or greater than an axial length (h) of the tip (64).

11. The gas turbine (10) according to any of the preceding claims, wherein the flow guiding elements (50) are configured such that a ratio h/P is equal to or greater than 0.75,

where:

h is an axial length of a tip (64) of an individual flow guiding element (50), and

P is a pitch between circumferentially neighboring flow guiding elements (50).

12. The gas turbine (10) according to any of the preceding claims, wherein the number of flow guiding elements (50) is an integral multiple of the number of diffuser struts (40).

13. The gas turbine (10) according to any of the preceding claims, wherein the flow guiding elements (50) are positioned with an angular offset in relation to the diffuser struts (40), such that each diffuser strut (40) has a

circumferential position between two circumferentially neighboring flow guiding elements (50).

14. The gas turbine (10) according to claim 13,

wherein the angular offset is in the range of 25 to 75% of an angular distance between circumferentially neighboring flow guiding elements (50).

15. The gas turbine (10) according to any of the preceding claims, wherein at least one of the flow guiding elements (50), as seen in top view, extends in a length direction along a straight line from a leading edge (52) to a trailing edge (54) thereof, the straight line being inclined in relation to the diffuser axis at angle (a) lying in the range of -25 to +25 degrees.

16. The gas turbine (10) according to any of the preceding claims, wherein each flow guiding element (50) comprises a pressure side (56) and a suction side (58) extending from a leading edge (52) to a trailing edge (54),

wherein the leading edge (52) is aerodynamically biased toward a predicted direction of flow (Fi) at the base load operation of the gas turbine.

17. The gas turbine (10) according to claim 16,

wherein a first intersection (66) of the leading edge (52) with the pressure side (56) is located forward of a second intersection (68) of the leading edge (52) with the suction side (58), as seen along a length direction of the flow guiding element (50), and

wherein the first intersection (66) comprises a sharp edge.

18. The gas turbine (10) according to claim 17, wherein the leading edge (52) includes a convex curved portion (92) between the first intersection (66) and the second intersection (68).

19. The gas turbine (10) according to claim 17, wherein the leading edge (52) includes a straight portion (96) between the first intersection (66) and the second intersection (68).

20. The gas turbine (10) according to any of claims 17 to 19, wherein an angle defined by the first intersection (66) is lesser than an angle defined by the second intersection (68).

Description:
GAS TURBINE EXHAUST DIFFUSER HAVING FLOW GUIDING

ELEMENTS

CROSS-REFERENCE TO RELATED APPLICATIONS

[0001] The present application claims priority to the U.S. Provisional Application No. 62/538,972, filed on July 31, 2017, and the U.S Provisional Application No. 62/549,603, filed on August 24, 2017, both of which are herein incorporated by reference in their entirety.

BACKGROUND 1. Field

[0002] The present invention relates to gas turbines, and in particular to a gas turbine exhaust diffuser having flow guiding elements.

2. Description of the Related Art

[0003] An axial flow turbomachine, such as a gas turbine engine, typically includes a compressor section for compressing air, a combustor section for mixing the compressed air with fuel and igniting the mixture to form a hot working medium fluid, a turbine section for extracting power from the working medium fluid, and an exhaust diffuser located downstream of a last turbine stage for recovering static pressure of the turbine exhaust flow.

[0004] An exhaust diffuser may serve to reduce the losses associated to the momentum of the flow exiting the turbine section, by reducing the high velocity of the flow leaving the last stage of the turbine blading to a moderate level. The reduction in velocity decreases the kinetic energy of the flow and increases the static pressure of the flow passing through the diffuser, before the flow leaves the exhaust system or enters a heat recovery steam generator.

[0005] As the rate of diffusion per unit length that can be achieved in such a diffuser is limited by the physics of the boundary layer flow in the diffuser, typical diffusers require a large amount of space, thus adding to the size and the cost of the plant arrangement. This may call for substantial efforts to increase the possible rate of diffusion in order to shorten the diffuser without compromising the effectiveness.

[0006] In typical heavy-duty gas turbine arrangements, one of the bearings supporting the rotor needs to be placed in the entry region of the diffuser immediately downstream of the turbine section. This requires the placement of supporting struts crossing the diffuser flow path in a region of relatively high velocity to transmit the bearing loads to the external engine supports and ultimately to the baseplate of the engine. The struts may desirably have an aerodynamic profile in order to minimize the drag of the struts and the associated loss in total pressure.

[0007] The direction of the flow entering the struts depends on the load set point of the engine. Typically, the struts are optimized for the flow direction at base load, which is close to the axial direction, i.e. parallel to the engine center line. The angle between the main flow direction and the engine center line changes towards the rotating direction of the turbine, when the engine is operated at part load conditions.

[0008] Due to increasing demand to extend the operating regime of gas turbines for power generation towards very low loads and to reduce the emissions under such operating conditions, the required operating range of the turbine diffusers has increased over the past decades. This results in a need to enable the diffuser to operate with high flow angles and flow incidence on the struts.

SUMMARY

[0009] Briefly, aspects of the present invention are directed to a gas turbine having flow guiding elements in an exhaust diffuser of the gas turbine.

[0010] According to an aspect of the present invention, a gas turbine is provided. The gas turbine comprises a turbine section, and a diffuser located downstream of a row of turbine blades of the turbine section. The diffuser comprises an annular duct extending axially along a diffuser axis. The annular duct is delimited radially by an outer wall and an inner wall which respectively define outer and inner boundaries of an exhaust flowpath. A plurality of diffuser struts are circumferentially distributed within the annular duct. Each diffuser strut extends from the outer wall to the inner wall. A plurality of flow guiding elements are circumferentially distributed within the annular duct. The flow guiding elements are positioned on the inner wall and extend radially therefrom into the exhaust flowpath. The flow guiding elements have an axial location downstream of the row of turbine blades and upstream of the diffuser struts. The number of flow guiding elements is larger than the number of diffuser struts.

BRIEF DESCRIPTION OF THE DRAWINGS

[0011] The invention is shown in more detail by help of figures. The figures show preferred configurations and do not limit the scope of the invention.

[0012] FIG. 1 is a schematic diagram of a gas turbine, wherein embodiments of the present invention may be employed;

[0013] FIG. 2 is a schematic axial end view of a diffuser section, looking in the direction of flow, according to an embodiment of the present invention;

[0014] FIG. 3 is a radial top view of a portion of the diffuser section, looking radially inward toward a diffuser hub;

[0015] FIG. 4 is a meridional view of a portion of the diffuser section;

[0016] FIG. 5 is a schematic diagram illustrating an arrangement of a flow guiding element in relation to a flow separation during operation;

[0017] FIG. 6, 7 and 8 show radial top views of flow guiding elements having various configurations of biased leading shapes, according to a further development of the present invention.

DETAILED DESCRIPTION

[0018] In the following detailed description of the preferred embodiments, reference is made to the accompanying drawings that form a part hereof, and in which is shown by way of illustration, and not by way of limitation, a specific embodiment in which the invention may be practiced. It is to be understood that other embodiments may be utilized and that changes may be made without departing from the spirit and scope of the present invention.

[0019] In the present description, a stated range is understood to include the boundary values of the range. That is, a range of X to Y is understood to include the values of X and Y.

[0020] Referring to FIG. 1, a gas turbine engine 10 (or simply "gas turbine") generally includes a compressor section 12, a combustor section 16, a turbine section 24 and an exhaust diffuser 30. In operation, the compressor section 12 inducts ambient air 14 and compresses it. The compressed air from the compressor section 12 enters one or more combustors in the combustor section 16. In the combustor section 16, the compressed air is mixed with a fuel 18, and the air-fuel mixture is burned in the combustors to form a hot gas 20, which forms a working medium fluid. The hot gas 20 is routed to the turbine section 24 where it is expanded through one or more turbine stages, each turbine stage comprising a row of stationary vanes followed by a row of rotating blades. The expansion of the hot gas is used to generate power that can drive a turbine rotor shaft 22. The expanded gas exiting the turbine section 24 is exhausted from the engine 10 via the exhaust diffuser 30 which is located downstream of a last turbine stage.

[0021] Referring to FIG. 2, the diffuser 30 is a stationary component comprising an annular duct 34 defining an exhaust flowpath and extending axially along a diffuser axis 32, which may be coaxial with the turbine rotor. The annular duct 34 is delimited radially by an outer wall 36 forming an outer flowpath boundary and an inner wall 38 forming an inner flowpath boundary. The outer wall 36 may be formed, for example, by a casing. The inner wall 38 may be formed by a hub. The annular duct 34 is typically configured to be divergent, being conical about the axis 32, which may serve to reduce the speed of the exhaust flow and thus increase the pressure difference of the exhaust gas expanding across the last stage of the turbine section 24. The diffuser 30 further includes supporting structures 40, referred to as struts, which are distributed circumferentially within the annular duct 34. Each strut 40 extends from the outer wall 36 to the inner wall 38, and may further extend through the outer wall 36 and the inner wall 38. To provide protection from high exhaust flowpath temperatures, each strut 40 may be provided with a protective shield that forms an outer surface of the strut 40. As shown in FIG. 3, the outer surface of each strut 40 may have an aerodynamic shape, comprising a leading edge 42 and a trailing edge 44.

[0022] As stated above, the direction of the flow entering the struts 40 depends on the load set point of the gas turbine engine. The orientation of the struts 40 may be optimized for the flow direction at base load operating conditions, which is close to the axial direction, i.e. parallel to the axis 32. When the engine is operated at part load conditions, the angle between the main flow direction and the axial direction changes towards the direction of rotation of the turbine rotor.

[0023] The present inventors have recognized that the high swirl or angle (in relation to the diffuser axis/ turbine rotor axis) of the flow exiting the turbine section at part load creates a radial gradient of the velocity profile with even further reduced velocity towards the inner wall of the flow path, which may result in flow separation at the inner wall immediately downstream of the last turbine stage. This flow separation may cause unsteady loads on the diffuser structure causing premature failure.

[0024] Embodiments of the present invention address at least some of the above- mentioned technical problems. Referring to FIG. 2-5, a diffuser 30 according to the illustrated embodiments comprises a plurality of flow guiding elements 50 circumferentially distributed within the annular duct 34. The flow guiding elements 50 are positioned on the inner wall 38 and extend radially from the inner wall 38 into the exhaust flowpath. The flow guiding elements 50 have an axial location which is downstream of a row of last stage turbine blades 26 and upstream of the diffuser struts 40. The embodiments illustrated herein are particularly effective in reducing the swirl of the flow incident on the struts 40 near the inner wall 38 of the exhaust flowpath. In addition, the aerodynamic turning of the flow guiding elements 50 serves to create stream-wise vortices that transport high momentum fluid from the main flow into the boundary layer, thus increasing the momentum of the boundary layer flow and further reducing the size of flow separation region at the inner wall 38 near the struts 40. To this end, the number of flow guiding elements 50 is larger than the number of struts 40. [0025] As shown in FIG. 3, each flow guiding element 50, as seen in top view, extends in a length direction along a straight line from a leading edge 52 to a trailing edge 54. The straight line may be parallel to the diffuser axis 32 or may be inclined thereto. The angle of inclination a may lie, for example, in the range of -25 to +25 degrees, especially in the range of -10 to +10 degrees. Herein, a positive (+) angle refers to an inclination in a direction U of rotation of the turbine rotor, while a negative (-) angle refers to an inclination opposite to the direction U of rotation of the turbine rotor. In the example of FIG. 3, a positive (+) angle of inclination is illustrated.

[0026] In the illustrated embodiment, the flow guiding elements 50 are axially located in a region in which flow is already separated from the inner wall 38 downstream of the last stage turbine blades 26, during operation of the gas turbine, in particular during a low power or part load operation. In one embodiment, to determine an axial location of the flow guiding elements 50, a predicted flow separation region 70 (see FIG. 5) may be determined, for example, using simulation tools, which may, for instance, be based on computational fluid dynamics (CFD) analyses.

[0027] The radial extension of the flow guiding elements 50 is smaller than the local distance between the outer wall 36 and the inner wall 38 of the annular duct 34. For the flow guiding elements 50 to be particularly effective, it is desirable that the radial extension of the flow guiding elements 50 is sufficient to penetrate through the separation region 70 downstream of the last stage turbine blades 26 and to reach into the main flow. Accordingly, in one embodiment, the local radial height r of a flow guiding element 50 is be greater than a predicted local radial height s of the flow separation region 70 over the inner wall 38, at a specified axial location. By extending beyond the separation region present at part load conditions, the flow guiding elements 50 are effective to turn the exhaust flow in the inner wall region towards the axial direction, thus reducing the aerodynamic load of the diffuser struts 40 and the risk of flow separation and mechanical excitation of the diffuser structure.

[0028] Referring to FIG. 4 and 5, each flow guiding element 50 has a base 62 adjoining the inner wall 38 and a tip 64 opposite to the base 62. The local radial height r of a flow guiding element 50 may be defined as a height of the tip 64 of the flow guiding element 50, at a specified axial location, as measured in a radial direction from the inner wall 38. The local radial height r of an individual flow guiding element 50 may vary between the leading edge 52 and the trailing edge 54 of the flow guiding element 50, as shown in FIG. 4 and 5. The local radial height s of the flow separation region 70 may be defined as a height of the flow separation region 70, at a specified axial location, as measured in a radial direction from the inner wall 38.

[0029] In some engine configurations, a high radial extension of the flow guiding elements 50 may cause aerodynamic drag and a risk of mechanical failure of the flow guiding elements 50. Accordingly, in one embodiment, the local radial height r of the flow guiding elements 50 may be designed to be less than or equal to 20% above the predicted local radial height s of the flow separation region 70 over the inner wall 38, at the specified axial location. Based on a different consideration, the local radial height r of the flow guiding elements 50 may be designed to lie in the range of 10 to 25%) of a local radial distance R between the outer wall 36 and the inner wall 38, at a specified axial location. The above range may also ensure an adequate penetration through the flow separation region during part load operation while reducing the risk of mechanical failure of the flow guiding elements 50. In the shown embodiment, a local radial height at the leading edge 52 of the flow guiding element 50 is lesser than a local radial height at the trailing edge 54 of the flow guiding element 50.

[0030] In order to maintain mechanical stability, it may be desirable that the axial length li of the base 62 of the flow guiding element 50 is equal to or greater than the axial length 1 2 of the tip 64 of the flow guiding element 50. In the shown embodiment, which is a non-limiting example configuration, the leading edge 52 of the flow guiding element 50 is inclined toward a downstream direction, while the trailing edge 54 of the flow guiding element 50 is inclined toward an upstream direction. The axial length li of the base 62 is, in this case, greater than the axial length I2 of the tip 64.

[0031] As stated above, the flow guiding elements 50 are desirably placed at an axial location where the flow has already separated downstream of the last stage turbine blades 26. Referring to FIG. 5, the radial height of the separation region 70 generally increases rapidly in the stream-wise or axial direction. Moving the flow guiding elements too far axially downstream may thereby increase the required radial height of the flow guiding elements 50 significantly, in order for the flow guiding elements 50 to penetrate through the flow separation region 70. The present inventors recognize that an increased radial height of the flow guiding elements 50 may increase the risk of mechanical failure as well as the aerodynamic drag of the flow guiding elements 50 at base load operating conditions. The present inventors further determined that moving the flow guiding elements 50 too far axially upstream may increase the risk of unfavorable excitation of the last stage turbine blades 26 by the bow waves penetrating upstream of the flow guiding elements 50. In accordance with various embodiments of the present invention, the axial location of the flow guiding elements 50 may be determined to address the aforementioned conflicting requirements.

[0032] In the illustrated embodiment, as shown in FIG. 4, the flow guiding elements 50 are axially located closer to an axial position of the trailing edge 28 of the last stage turbine blades 26 than to an axial position of the leading edge 42 of the struts 40. A first axial distance di may be defined between an axial position of the leading edge 52 of the flow guiding elements 50 and the axial position of the trailing edge 28 of the last stage turbine blades 26. A second axial distance d 2 may be defined between the axial position of the trailing edge 28 of the last stage turbine blades 26 and the axial position of the leading edge 42 of the struts 40. In one embodiment, the axial placement of the flow guiding elements 50 may be determined such that the distance di lies in the range of 5 to 25% of the distance d 2 . In this context, the axial distances di and d 2 are measured at the hub-side, i.e., at the inner wall 38. Alternately or additionally, the axial placement of the flow guiding elements 50 may be determined in terms of a hub side chord length of the last stage turbine blades 26, such that the distance di lies in the range of 5 to 25% of the hub side chord-length of the last stage turbine blades 26. The hub side chord length of the blades 26 may be defined as a straight line distance between the leading edge and the trailing edge of the blade airfoils at the hub or inner diameter end of the blade airfoils.

[0033] It has been determined that in some engine configurations, at part load or low load operating conditions, the vortices forming at the tips of the individual flow guiding elements 50 tend to propagate downstream in a direction close to the direction of the main flow that has very high swirl. In such a configuration, the circumferential spacing between neighboring flow guiding elements 50 may be desirably determined such that the vortex trajectories of a flow guiding element 50 do not interfere with a neighboring flow guiding element 50. Accordingly, in one embodiment, the flow guiding elements 50 may be configured such that the ratio fP is greater than or equal to 0.75, where h is the axial length of the tip 64 of the flow guiding element 50 (see FIG. 5), and P is the pitch between circumferentially neighboring flow guiding elements 50 (see FIG. 3).

[0034] In some engine configurations, the flow propagating downstream of the individual flow guiding elements 50 may interact with the diffuser struts 40. A further improvement may be achieved in this case by ensuring that the number of the flow guiding elements 50 is an integral multiple of the number of struts 40. Such a design may allow an arrangement of the individual flow guiding elements 50 such that their circumferential position relative to the downstream struts 40 creates a pattern, which is periodically repeated along the circumference with the angular pitch of the struts 40, reducing the risk of unintended asymmetries of the flow approaching the struts 40. In the shown example (see FIG. 2), which is non-limiting, the ratio of the number of flow guiding elements to the number of struts is seven. In general, this ratio may assume any integer value equal to or greater than two. Additionally, or alternately, the position of the individual flow guiding elements 50 may be determined such that the wakes created by the individual flow guiding elements 50 do not impinge on the struts 40 at base load conditions. Accordingly, in one embodiment, as shown in FIG. 2 and 3, the flow guiding elements 50 may be positioned with an angular offset in relation to the struts 40, such that each strut 40 has an angular position between two circumferentially neighboring flow guiding elements 50. In particular, the angular offset may lie in the range of 25 to 75%, in particular about 50 %, of an angular distance between two neighboring flow guiding elements 50.

[0035] In a further development, the flow guiding elements may be designed such that the leading edge of one or more of the flow guiding elements is aerodynamically biased toward incident flow at high power or base load operating conditions. The flow guiding elements may be thereby configured to create little impact (i.e., minimum losses) at high power or base load operating conditions, while at low power or part load operating conditions, the high angle of incidence would cause vortex shedding, due to the biased shape of the leading edge.

[0036] Example embodiments of such flow guiding elements are illustrated in FIG. 6, 7 and 8. Herein, the arrows Fi indicate the dominant direction of flow during a high power or base load operating condition, while the arrows F 2 indicate the dominant direction of flow during a low power or part load operating condition. The directions Fi and F 2 may be dependent on engine operating parameters, including but not limited to, ambient outside temperature, power output, injection of water or water vapor or implementation of other cooling devices to reduce inlet flow temperature, presence of a heat recovery steam generator (FIRSG) aft of the exhaust diffuser 30, among others. The directions Fi and F 2 may be predicted for a given engine configuration, for example, based on CFD analyses. In general, the flow direction Fi at high power or base load conditions is close to parallel to the engine centerline (diffuser axis 32), or inclined slightly thereto, particularly in a direction opposite to the direction U of rotation of the turbine rotor. At low power or part load operating conditions, the flow direction F 2 is inclined toward the direction U of rotation of the turbine rotor.

[0037] As shown, each flow guiding element 50 comprises a pressure side 56 and a suction side 58, which extend from the leading edge 52 to the trailing edge 54. In this context, the pressure side 56 and suction side 58 are specified in reference to the rotation direction U, such that the pressure side 56 is positioned aft of the suction side 58 in relation to the rotation direction U. As shown, the pressure side 56 and the suction side 58 may extend parallel to the length direction of the flow guiding element 50. In the illustrated embodiments of FIG. 6-8, the length direction of the flow guiding element 50 may be understood to be parallel to the diffuser axis 32, it being understood that the underlying concepts could be also applied to a configuration where the flow guiding elements 50 are inclined to the diffuser axis 32 (e.g. as shown in FIG. 3). In each of the illustrated embodiments, the leading edge 52 of the flow guiding elements 50 is aerodynamically biased toward the flow direction Fi at high power or base load operating conditions. This may be achieved by designing the flow guiding element 50 such that a first intersection 66 of the leading edge 52 with the pressure side 56 is located forward of a second intersection 68 of the leading edge 52 with the suction side 58, as seen along the length direction of the flow guiding element 50. This ensures that the leading edge 52 is misaligned with the flow direction F 2 at low power or part load operating conditions. The first intersection 66 includes a sharp edge, which is configured to produce vortices 80 on the pressure side 56 of the flow guiding elements 50 at low power or part load operating conditions. The formation of vortices 80 on the pressure side 56 further reduces flow separation at the inner wall of the flowpath at part load operation, thereby improving performance at part load operation with minimum losses at base load operation.

[0038] In the embodiment shown in FIG. 6, the leading edge 52 includes a convex curved portion 92 between the first intersection 66 and the second intersection 68. The convex curved portion 92 may include, for example, a smooth elliptical shape to provide low loss at base load operation. The convex curved portion 92 extends seamlessly (i.e., smoothly) from the intersection 68 and is abruptly truncated at the intersection 66 to form a sharp edge, which aids formation of vortices 80 on the pressure side 56 at part load operation. In the embodiment shown in FIG. 7, the leading edge 52 includes a straight portion 96 connecting the first intersection 66 and the second intersection 68. In yet another embodiment as shown in FIG. 8, the leading edge 52 may include multiple segments 94, 96 between the first intersection 66 and the second intersection 68. The segments 94, 96 may have straight or curved profiles. In each embodiment, the angle defined by the first intersection 66 is lesser than the angle defined by the second intersection 68, to produce a relatively sharp edge at the first intersection 66.

[0039] While specific embodiments have been described in detail, those with ordinary skill in the art will appreciate that various modifications and alternative to those details could be developed in light of the overall teachings of the disclosure. Accordingly, the particular arrangements disclosed are meant to be illustrative only and not limiting as to the scope of the invention, which is to be given the full breadth of the appended claims, and any and all equivalents thereof.