ROTATING TURBOMACHINE COMPONENT HAVING A TIP LEAKAGE FLOW GUIDE

Abstract

A rotating turbomachine component includes a base portion and an airfoil portion extending from the base portion. The airfoil portion includes a first end connected to the base portion and a tip end portion that is cantilevered from the base portion. A tip leakage flow guide is provided at the tip end portion of the airfoil portion. The tip leakage flow guide includes one or more turning vane members configured and disposed to guide a leakage flow from the tip end portion at a flow angle that substantially coincides with a flow angle of gases flowing downstream from the rotating turbomachine component.

Description

BACKGROUND OF THE INVENTION

The subject matter disclosed herein relates to the art of turbomachines and, more particularly, to a rotating turbomachine component having a tip leakage flow guide.

Many turbomachines include a compressor portion linked to a turbine portion through a common compressor/turbine shaft or rotor and a combustor assembly. The compressor portion guides a compressed air flow through a number of sequential stages toward the combustor assembly. In the combustor assembly, the compressed air flow mixes with a fuel to form a combustible mixture. The combustible mixture is combusted in the combustor assembly to form hot gases. The hot gases are guided to the turbine portion through a transition piece. The hot gases expand through the turbine rotating turbine blades to create work that is output, for example, to power a generator, a pump, or to provide power to a vehicle. In addition to providing compressed air for combustion, a portion of the compressed airflow is passed through the turbine portion for cooling purposes.

In some cases, the hot gases expanding through the turbine portion leak or pass over tip end portions of the turbine blades. In order to reduce leakage, manufactures maintain tight clearances between the tip end portions and stationary components of the turbomachine. Generally, seals are provided on the stationary component or turbine shroud. While effective, existing seals still allow a portion of the hot gases or leakage gases to pass over the tip end portion. The tight clearance established by the seal causes the leakage gases to exit at an angle that is generally parallel to an axis defined by a turbomachine rotor. In contrast, hot gases passing along the gas path exit the rotor blades at an angle. Interactions between the leakage gases and the hot gases flowing along the gas path create localized pressure drops that have a negative impact on turbomachine performance.

BRIEF DESCRIPTION OF THE INVENTION

According to one aspect of the exemplary embodiment, a rotating turbomachine component includes a base portion and an airfoil portion extending from the base portion. The airfoil portion includes a base portion and a tip end portion that is cantilevered from the base portion. A tip leakage flow guide is provided at the tip end portion of the airfoil portion. The tip leakage flow guide includes one or more turning vane members configured and disposed to guide a leakage flow from the tip end portion at a flow angle that substantially coincides with a flow angle of gases flowing downstream from the rotating turbomachine component.

According to another aspect of the exemplary embodiment, a method of operating a turbomachine includes passing hot gases from a combustor assembly toward a plurality of buckets, guiding the hot gases onto the plurality of buckets, directing the hot gases downstream relative to the plurality of buckets along a gas path at a first flow angle, passing a portion of the hot gases over a tip end portion of the plurality of buckets at a second flow angle that is distinct from the first flow angle, and guiding the portion of the hot gases from the tip end portion of the plurality of buckets at a third flow angle that substantially coincides with the first angle.

According to yet another aspect of the exemplary embodiment, a turbomachine includes a compressor portion, a combustor assembly fluidly connecting the compressor portion and a turbine portion mechanically linked to the compressor portion and fluidly connected to the combustor assembly. The turbine portion includes a rotating component having a base portion and an airfoil portion extending from the base portion. The airfoil portion includes a first end connected to the base portion and a tip end portion that is cantilevered from the base portion. A tip leakage flow guide is provided at the tip end portion of the airfoil portion. The tip leakage flow guide includes one or more turning vane members configured and disposed to guide a leakage flow from the tip end portion at a flow angle that substantially coincides with a flow angle of gases flowing downstream from the rotating turbomachine component. A turning vane support member is positioned at the tip end portion. The turning vane support member includes an upstream end and a downstream end. The one or more turning vane members project outward from the turning vane support member

These and other advantages and features will become more apparent from the following description taken in conjunction with the drawings.

BRIEF DESCRIPTION OF DRAWINGS

The subject matter, which is regarded as the invention, is particularly pointed out and distinctly claimed in the claims at the conclusion of the specification. The foregoing and other features, and advantages of the invention are apparent from the following detailed description taken in conjunction with the accompanying drawings in which:

FIG. 1 is a schematic view of a turbomachine including a tip leakage flow guide in accordance with an exemplary embodiment;

FIG. 2 is a partial cross-sectional view of the turbomachine of FIG. 1;

FIG. 3 is a detail view of a rotating component of the turbomachine of FIG. 1 including a tip leakage flow guide in accordance with an exemplary embodiment;

FIG. 4 is a perspective view of the tip leakage flow guide of FIG. 3 having a plurality of turning vane members in accordance with one aspect of the exemplary embodiment;

FIG. 5 is a perspective view of the tip leakage flow guide of FIG. 3 having a plurality of turning vane members in accordance with another aspect of the exemplary embodiment; and

FIG. 6 is a perspective view of the tip leakage flow guide of FIG. 3 having a plurality of turning vane members in accordance with still another aspect of the exemplary embodiment.

The detailed description explains embodiments of the invention, together with advantages and features, by way of example with reference to the drawings.

DETAILED DESCRIPTION OF THE INVENTION

With reference to FIGS. 1 and 2, a turbomachine constructed in accordance with an exemplary embodiment is indicated generally at 2. Turbomachine 2 includes a compressor portion 4 operatively connected to a turbine portion 6. A combustor assembly 8 is fluidly connected to compressor portion 4 and turbine portion 6. Combustor assembly 8 is formed from a plurality of circumferentially spaced combustors, one of which is indicated at 10. Of course it should be understood that combustor assembly 8 could include other arrangements of combustors. Compressor portion 4 is also linked to turbine portion 6 through a common compressor/turbine shaft 12. With this arrangement, compressor portion 4 delivers compressed air to combustor assembly 8. The compressed air mixes with a combustible fluid to form a combustible mixture. The combustible mixture is combusted in combustor 10 to form products of combustion that are delivered to turbine portion 6 through a transition piece (not shown). The products of combustion expand along a gas path 18 of turbine portion 6 to power, for example, a generator, a pump, or a vehicle or the like (also not shown).

In the exemplary embodiment shown, turbine portion 6 includes a housing 19 that encases a first, stage 20 and a second stage 21 that define gas path 18. First stage 20 includes a plurality of first stage stators or nozzles, one of which is indicated at 30, supported to turbine housing 19 through a nozzle platform 31. First stage 20 also includes a plurality of first stage buckets or blades, one of which is indicated at 32, mounted to a first stage rotor wheel 34. Blades 32 are spaced from a stationary shroud member 35. Blades 32 include a base portion 38 and an airfoil portion 40. Airfoil portion 40 includes a first end 42 coupled to base portion 38 and a second end or tip end portion 44 that is spaced from stationary shroud member 35. Second stage 21 includes a plurality of second stage stators or nozzles, one of which is indicated at 48 supported to turbine housing 19 through a nozzle platform 49. Second stage 21 also includes a plurality of second stage buckets or blades, one of which is indicated at 50. At this point it should be understood that the number of stages in turbine portion 6 could vary.

In accordance with an exemplary embodiment, turbomachine 2 includes a tip leakage flow guide 60 that conditions tip leakage flow passing over tip portions of blades 32. As best shown in FIG. 3, tip leakage flow guide 60 includes a turning vane support member 64 mounted to tip end portion 44 of blade 32. Turning vane support member 64 includes an upstream end 66 that extends to a downstream end 68 through a substantially planar surface 70. A seal element 74 extends from substantially planar surface 70 into a pocket (not separately labeled) of stationary shroud member 35. Seal element 74 limits flow passing from gas path 18 across tip end portion 44 of blade 36. However, while reduced, some leakage flow does flow over tip end portion 44 despite the presence of seal element 74. In order to reduce losses associated with the leakage flow, one or more turning vane members 80 are positioned on turning vane support member 64. In the exemplary aspect shown, turning vane member 80 is arranged adjacent to downstream end 68. Turning vane member 80 alters a flow path of the leakage flow.

Combustion gases flow along gas path 18 and pass over nozzles 30 and are guided toward blades 32. A first or main flow 85 passes over blades 32 and a second or leakage flow 88 passes over tip end portion 44 along gas path 18. Main flow 85 flows at a first flow angle as a result of interactions with blade 36. Leakage flow 88 flows at a second flow angle, that is distinct from the first flow angle, and which runs generally parallel to shaft 12. Turning vane member 80 is configured to condition or turn leakage flow 88 exiting tip end portion 44 to create a turned flow 91 that returns to gas path 18 at a third flow angle that substantially coincides with the first flow angle of main flow 85 flowing downstream from blades 32. By matching the third flow angle with the first flow angle, undesirable interactions between turned flow 91 and main flow 85 are reduced. In this manner, turning vane member 80 reduces losses within turbine portion 6 associated with pressure variations along gas path 18 resulting from undesirable interactions between leakage flow 88 and the main flow 85. In the event that nozzles 30 form part of a last stage (not separately labeled) of turbine portion 6, turning vane 80 may be configured to guide the leakage flow gases at an angle that generally corresponds to the flow angle of gases flowing downstream toward and along a radial diffusion section (not shown) of turbine portion 6 so as to enhance pressure recovery.

In accordance with one aspect of the exemplary embodiment illustrated in FIG. 4, turning vane member 80 takes the form of a plurality of substantially linear vane members 97. Each vane member 97 includes a first end 99 and a second end 100. Second end 100 is off-set relative to first end 99 such that vane members 97 are angled relative to, for example, shaft 12. More specifically, vane members 97 are angled so as to generally correspond to an airfoil profile 102 of airfoil portion 40. In accordance with one aspect of the exemplary embodiment, the angle of vane members 97 is substantially equal to or ±30° of a trailing edge angle θ of airfoil profile 102. FIG. 5 illustrates turning vanes 106 in accordance with another aspect of the exemplary embodiment. Turing vanes 106 take the form of a plurality of curvilinear vane members 110 having first and second curvilinear surfaces 112 and 113. In a manner similar to that described above, vane members 110 are angled so as to generally correspond to an airfoil profile 102 of airfoil portion 40. In accordance with one aspect of the exemplary embodiment, the angle of vane members 110 is substantially equal to or ±30° of a trailing edge angle θ of airfoil profile 102. FIG. 6 illustrates turning vanes 117 in accordance with yet another aspect of the exemplary embodiment. Turning vanes 117 take the form of complex geometrical vane members 121. Complex geometrical vane members 121 include a first vane member 123 and a second vane member 124. First vane member 123 includes a first end section 126 that extends to a second end section 127. Second vane member 124 includes a first end portion 129 that extends from second end section 127 of first vane member 123 to a second end portion 130. Second end portion 130 is off-set relative to first end section 126 of first vane member 123 and is angled so as to generally correspond to an airfoil profile 102 of airfoil portion 40. In accordance with one aspect of the exemplary embodiment, the angle of second end portion 130 is substantially equal to or ±30° of a trailing edge angle θ of airfoil profile 102. Regardless of form, the turning vanes condition the leakage flow to pass back into the gas path at an angle the substantially coincides with the main flow to reduce undesirable interactions.

At this point it should be understood that the exemplary embodiments provide a system for redirecting tip leakage flow back into the gas path to reduce undesirable interactions with the main flow. Reducing undesirable interactions with the main flow leads to a reduction in pressure losses that may detract from turbine performance. It should also be understood that while shown in connection with a gas turbomachine, the exemplary embodiments could also be employed in a steam turbomachine.

While the invention has been described in detail in connection with only a limited number of embodiments, it should be readily understood that the invention is not limited to such disclosed embodiments. Rather, the invention can be modified to incorporate any number of variations, alterations, substitutions or equivalent arrangements not heretofore described, but which are commensurate with the spirit and scope of the invention. Additionally, while various embodiments of the invention have been described, it is to be understood that aspects of the invention may include only some of the described embodiments. Accordingly, the invention is not to be seen as limited by the foregoing description, but is only limited by the scope of the appended claims.

Claims

1. A rotating turbomachine component comprising:

a base portion;

an airfoil portion extending from the base portion, the airfoil portion including a first end connected to the base portion and a tip end portion that is cantilevered from the base portion; and

a tip leakage flow guide provided at the tip end portion of the airfoil portion, the tip leakage flow guide including one or more turning vane members configured and disposed to guide a leakage flow from the tip end portion at a flow angle that substantially coincides with a flow angle of gases flowing downstream from the rotating turbomachine component.

2. The rotating turbomachine component according to claim 1, further comprising: a turning vane support member positioned at the tip end portion, the turning vane support member having an upstream end, and a downstream end, each of the upstream end and the downstream end projecting beyond the tip end portion, the one or more turning vane members projecting outward from the turning vane support member.

3. The rotating turbomachine component according to claim 2, wherein the tip leakage flow guide is arranged at the downstream end of the turning vane support member.

4. The rotating turbomachine component according to claim 1, wherein the one or more turning vane members comprise a plurality of substantially linear vane members extending across the tip end portion, each of the plurality of substantially linear vane members including a first end and a second end, the second end being off-set relative to the first end.

5. The rotating turbomachine component according to claim 1, wherein the one or more turning vane members comprise a plurality of curvilinear vane members extending across the tip end portion.

6. The rotating turbomachine component according to claim 1, wherein the one or more turning vane members comprise a plurality of complex geometrical vane members extending across the tip end portion.

7. The rotating turbomachine component according to claim 6, wherein each of the plurality of complex geometrical vane members including a first vane member having a first end section that extends to a second end section, and a second vane member having a first end portion that extends from the second end section of the first vane member to a second end portion, the second end portion being off-set relative to the first end section.

8. The rotating turbomachine component according to claim 1, wherein the one or more turning vane members are arranged at an angle that generally corresponds to an airfoil profile of the airfoil portion.

9. A method of operating a turbomachine comprising:

passing hot gases from a combustor assembly toward a plurality of buckets;

guiding the hot gases onto the plurality of buckets;

directing the hot gases downstream relative to the plurality of buckets along a gas path at a first flow angle;

passing a portion of the hot gases over a tip end portion of the plurality of buckets at a second flow angle that is distinct from the first flow angle; and

guiding the portion of the hot gases from the tip end portion of the plurality of buckets at a third flow angle that substantially coincides with the first flow angle.

10. The method of claim 9, wherein passing the portion of hot gases from the tip end portion includes guiding the portion of hot gases across one or more turning vane members arranged at the tip end portion.

11. The method of claim 10, wherein guiding the portion of hot gases across one or more turning vane members includes passing the portion of hot gases over a plurality of angled vane members.

12. The method of claim 10, wherein guiding the portion of hot gases across one or more turning vane members includes passing the portion of hot gases over a plurality of curvilinear vane members.

13. The method of claim 10, wherein guiding the portion of hot gases across one or more turning vane members includes passing the portion of hot gases at an angle that generally corresponds to an angle of an airfoil portion of each of the plurality of buckets.

14. A turbomachine comprising:

a compressor portion;

a combustor assembly fluidly connecting the compressor portion;

a turbine portion mechanically linked to the compressor portion and fluidly connected to the combustor assembly, the turbine portion including a rotating component having a base portion and an airfoil portion extending from the base portion, the airfoil portion including a first end connected to the base portion and a tip end portion that is cantilevered from the base portion;

a tip leakage flow guide provided at the tip end portion of the airfoil portion, the tip leakage flow guide including one or more turning vane members configured and disposed to guide a leakage flow from the tip end portion at a flow angle that substantially coincides with a flow angle of gases flowing downstream from the rotating component; and

a turning vane support member positioned at the tip end portion, the turning vane support member having an upstream end and a downstream end, the one or more turning vane members projecting outward from the turning vane support member.

15. The turbomachine according to claim 14, wherein the one or more turning vane members is arranged at an angle that generally corresponds to an airfoil profile of the airfoil portion.

16. The turbomachine according to claim 15, wherein the angle of the one or more turning vane members is within no more than about 30° of a trailing edge angle of the airfoil profile.

17. The turbomachine according to claim 14, wherein the one or more turning vane members comprise a plurality of substantially linear vane members extending across the tip end portion, each of the plurality of substantially linear vane members including a first end and a second end, the second end being off-set relative to the first end.

18. The turbomachine according to claim 13, wherein the one or more turning vane members comprise a plurality of curvilinear vane members extending across the tip end portion.

19. The turbomachine according to claim 13, wherein the one or more turning vane members comprise a plurality complex geometrical vane members extending across the tip end portion.

20. The turbomachine according to claim 19, wherein the plurality of complex geometrical vane members include a first vane member having a first end section that extends to a second end section, and a second vane member having a first end portion that extends from the second end section of the first vane member to a second end portion, the second end portion being off-set relative to the first end section.