CROSS REFERENCE TO RELATED APPLICATIONS

[0001] This application is the International Application of US National Stage Application 62/043,452, filed August 29, 2014 and claims the benefit thereof. All of the applications are incorporated by reference herein in their entirety.

BACKGROUND

1. Field

[0002] The present invention relates to turbine engines, and more specifically to a flow guiding tip shroud for a turbine blade.

2. Description of the Related Art

[0003] In an industrial gas turbine engine, hot compressed gas is produced. The hot gas flow is passed through a turbine and expands to produce mechanical work used to drive an electric generator for power production. The turbine generally includes multiple stages of stator vanes and rotor blades to convert the energy from the hot gas flow into mechanical energy that drives the rotor shaft of the engine. Turbine inlet temperature is limited by the material properties and cooling capabilities of the turbine parts.

[0004] A combustion system receives air from a compressor and raises it to a high energy level by mixing in fuel and burning the mixture, after which products of the combustor are expanded through the turbine.

[0005] Gas turbines are becoming larger, more efficient, and more robust. Large blades and vanes are being produced, especially in the hot section of the engine system. Of particular challenge is the last stage blade. Traditionally the last stage blade has been solid, tip shrouded and uncooled. This configuration has limitations as the blades require more robustness as the gas path diameters increase and the gas path temperatures increase.

[0006] A turbine blade is formed from a root portion coupled to a rotor disc and an airfoil that extends outwardly from a platform coupled to the root portion. The blade is ordinarily composed of a tip opposite the root section, a leading edge, and a trailing edge. The tip of a turbine blade often has a tip feature to reduce the size of the gap between stationary casing and rotating blades in the gas path of the turbine to prevent tip flow leakage. The tip leakage flow reduces the amount of torque generated by the turbine blades, however, the tip feature may mitigate the leakage as much as possible. Some turbine blades include tip shrouds, as shown in FIG 1 , attached to the blade tips. To reduce over-tip leakage, shrouded blades may include one or more circumferential knife edges for running tight tip gaps. The turbine tip shrouds are also used for the purpose of damping blade mechanical vibrations, particularly for blades having high aspect ratio such as those used in lower pressure turbine stages.

[0007] Some modern tip shrouds are scalloped, as opposed to a full coverage tip shroud, to reduce shroud weight and hence lower centrifugal pull loads. The material removed by scalloping is indicated by the scalloped region 62 in FIG 4. The removal of material by scalloping increases aerodynamic losses thereby reducing the stage efficiency.

[0008] In current assemblies, the rotating blade tip shroud and cavity configurations in large industrial gas turbines are regions of low performance. There are several drivers of aerodynamic loss in the turbine-shroud cavity configuration, which lowers the gas turbine's efficiency. One driver is the flow over the rotating blade tip seal. Tip seals are generally designed to choke the flow and consequently lead to high flow velocities in the turbine tip-shroud cavity. The mixing losses that occur downstream of the seal are high and contribute to a reduction in stage efficiency and power. Additional mixing losses occur when the flow through the tip cavity combines with the main flow and the two streams have different velocities.

[0009] There have been several previous configuration modifications to the tip shroud that have been attempted to reduce loss associated with flow over the shroud, such as is shown in FIG 3. Here, a wall 56 provides for the extraction of a flux to be accelerated by the shroud to be effected through an annular grid 58 along a diffuser 60.

SUMMARY

[0010] In one aspect of the present invention, a blade for a turbine engine comprises: an airfoil extending span-wise along a radial direction; and a shroud positioned along a tip of the airfoil extending generally along a circumferential direction, scalloped, across a pressure side and a suction side of the airfoil, the shroud comprising: an upstream edge and a downstream edge spaced apart from each other in an axial direction; a radially inner surface adjoining the tip of the airfoil and a radially outer surface opposite to the radially inner surface; a shroud ridge extending across the shroud along a portion of the radially outer surface, wherein the shroud ridge is substantially rectangular in shape; a plurality of bladelets attached across a span of the shroud ridge in a row, wherein the plurality of bladelets produce a plurality of passages with each passage in between each bladelet, wherein half-width passages lie along each of the extreme edges of the plurality of bladelets; an intershroud bridge that extends along a portion of the shroud ridge closing over at least one half-width passage along at least one end of the shroud ridge at an extreme of the plurality of bladelets, the intershroud configured to reduce fluid separation adjacent to the plurality of bladelets, the shroud constructed and arranged to minimize losses; and a seal extending radially outward from the outer surface of the shroud.

[0011] In another aspect of the present invention, a method for aligning an asymmetric flow field along a tip shroud with flow guiding features comprises: providing a blade for a turbine engine comprising: an airfoil extending span-wise along a radial direction; and a shroud positioned along a tip of the airfoil extending generally along a circumferential direction, scalloped, across the pressure side and suction side of the airfoil, the shroud comprising: an upstream edge and a downstream edge spaced apart from each other in an axial direction; a radially inner surface adjoining the tip of the airfoil and a radially outer surface; a shroud ridge extending across the shroud along a portion of the radially outer surface; and a plurality of bladelets attached across a span of the shroud ridge in a row, wherein the plurality of bladelets produce a plurality of passages with each passage in between each bladelet, wherein half-width passages lie along each of the extreme edges of the plurality of bladelets; extending an intershroud bridge along a portion of the shroud ridge closing over at least one half-width passage along at least one extreme end of plurality of bladelets on the shroud ridge, the intershroud bridge configured to reduce fluid separation adjacent to the plurality of bladelets, the shroud constructed and arranged to minimize losses; and extending a seal radially outward from the radially outer surface of the shroud.

[0012] These and other features, aspects and advantages of the present invention will become better understood with reference to the following drawings, description and claims.

BRIEF DESCRIPTION OF THE DRAWINGS

[0013] 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.

[0014] FIG 1 is a perspective view of a gas turbine engine with a row of shrouded turbine blades wherein embodiments of the present invention may be incorporated.

[0015] FIG 2 is an axial cross-section of a turbine airfoil tip shroud and cavity configuration.

[0016] FIG 3 is a perspective view of a modified tip shroud of the prior art.

[0017] FIG 4 is a perspective view of turbine airfoil tip shroud bladelets of the prior art.

[0018] FIG 5 is perspective view of a turbine airfoil shroud bladelets with half-width passages blocked in an exemplary embodiment of the present invention.

[0019] FIG 6 is axial view of a turbine shroud bladelets with non-uniform pitch/chord ratio in an exemplary embodiment of the present invention.

[0020] FIG 7 is an axial view of a turbine tip shroud scalloping in an exemplary embodiment of the present invention.

[0021] FIG 8 is a perspective view of a turbine tip shroud ridge with bladelets and passages.

[0022] FIG 9 is an axial cross sectional view of a bladelet and shroud ridge corners in an exemplary embodiment of the present invention.

[0023] FIG 10 is an axial view of an exemplary embodiment of the present invention.

[0024] FIG 1 1 is a top perspective view of a turbine blade and tip shroud of an exemplary embodiment of the present invention.

[0025] FIG 12 is an axial view of the tip shroud cavity configuration of an exemplary embodiment of the present invention along A-A in Fig 1 1.

[0026] FIG 13 is an axial view of the tip shroud cavity configuration of an exemplary embodiment of the present invention along A-A in Fig 1 1.

[0027] FIG 14 perspective view of the tip shroud cavity configuration of an exemplary embodiment of the present invention along B-B in Fig 1 1.

[0028] FIG 15 is a perspective view of a tip shroud and cavity configuration in an exemplary embodiment of the present invention.

[0029] FIG 16 is a detailed radial view of a tip shroud and cavity configuration in an exemplary embodiment of the present invention. [0030] FIG 17 is a circumferential view of a tip shroud and cavity configuration in an exemplary embodiment of the present invention.

DETAILED DESCRIPTION

[0031] In the following detailed description of the preferred embodiment, 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.

[0032] Broadly, an embodiment of the present invention provides a turbine blade includes an airfoil extending span-wise along a radial direction, and a shroud positioned along a tip of the airfoil extending generally along a circumferential direction, scalloped, across a pressure side and a suction side of the airfoil, the shroud includes an upstream edge and a downstream edge spaced apart axially. The shroud further includes a radially inner surface adjoining the tip of the airfoil and a radially outer surface opposite to the radially inner surface. A shroud ridge extends across the shroud along a portion of the radially outer surface. A plurality of bladelets is attached across a span of the shroud ridge. The plurality of bladelets produces a plurality of passages with each passage in between each bladelet. Half-width passages lie along each of an extreme edge of the plurality of bladelets. An intershroud bridge that extends along a portion of the shroud ridge closes over at least one half-width passage along at least one end of the shroud ridge at an extreme end of the plurality of bladelets. The shroud also include a seal extends radially outward from the radially outer surface of the shroud.

[0033] A gas turbine engine may comprise a compressor section, a combustor and a turbine section. The compressor section compresses ambient air. The combustor combines the compressed air with a fuel and ignites the mixture creating combustion products comprising hot gases that form a working fluid. The working fluid travels to the turbine section. Within the turbine section are circumferential alternating rows of vanes and blades, the blades being coupled to a rotor. Each pair of rows of vanes and blades forms a stage in the turbine section. The turbine section comprises a fixed turbine casing, which houses the vanes, blades and rotor.

[0034] Any leakage flow, or aerodynamic loss, that is not turned by the blades is lost work extraction, thus lowering the turbine efficiency. One area of concern is the flow over the rotating blade tip seal. The mixing losses that occur downstream of the seal are high and contribute to a reduction in stage efficiency and power. Additional mixing losses occur when the flow through the tip cavity combines with the main flow and the two streams have different velocities.

[0035] A reduction in loss by providing a more uniform flow that decreases the likelihood of flow coming up and over the blade and increases the likelihood of flow staying within the blade passage is desirable. Embodiments of the present invention provide a tip shroud configuration for a blade that may allow for the reduction in losses.

[0036] Increasing the flow area and enabling the leakage flow to turn into alignment with the main gas path flow may improve the power output and performance of the stage. Previous attempts to design a tip seal with an airfoil-like passage or passages in order to help direct the flow in the most efficient manner and reduce the mixing losses have not been totally successful due to the highly unsteady nature of the flow field in this region as is seen in Figure 3. As configured in Figure 4 the blade let profile and tip vortex losses cancel efficiency gains from reduced mixing losses. This is due to an asymmetric flow field through the bladelets which emanates primarily from the scalloped shroud.

[0037] Referring to FIG 1 , a portion of a turbine section of a gas turbine engine 56 is shown, which includes a row of turbine blades 10 wherein embodiments of the present invention may be incorporated. The blades 10 are circumferentially spaced apart from each other to define respective flow passages between adjacent blades 10, for channeling the working fluid. The blades 10 are rotatable about a rotation axis along a centerline 1 1 of the turbine engine 56. Each blade 10 is formed from an airfoil 12 extending span-wise in a radial direction in the turbine engine 56 from a rotor disc. The airfoil 12 includes a leading edge 14, a trailing edge 16, a pressure side 18, a suction side 20 on a side opposite to the pressure side 18, a tip 22 at a radially outer end of the airfoil 12, a platform 24 coupled to the airfoil 12 at a radially inner end of the airfoil 12 for supporting the airfoil 12 and for coupling the airfoil 12 to the rotor disc. The blade 10 may further include a shroud 26, referred to as a tip shroud, coupled to the tip 22 of the generally elongated airfoil 12. The platform 24 forms a radially inner end wall, while the shroud 26 forms a radially outer end wall of the blade 10.

[0038] The shroud 26 includes an upstream edge 28 and a downstream edge 30. A radially inner surface 32 adjoins the tip 22 of the airfoil 12. The shroud 26 includes a radially outer surface 34 opposite to the radially inner surface 32. The radially inner surface 32 and the radially outer surface 34 connect at the upstream edge 28 and the downstream edge 30.

[0039] FIG 2 and FIG 1 1 show the area around the tip shroud 26 and a cavity 36 configuration in a detailed view. The shroud 26 may extend along a circumferential direction. The shrouds 26 of adjacent blades 10 may adjoin in the circumferential direction to form a shroud ring. A seal 38 may be provided on the shroud 26, extending radially outward from the radially outer surface 34 of the shroud 26. The seal 38 may have a knife edge. The seal 38 may run a tight tip gap against a turbine component 40 of the turbine engine 56 reducing overtip leakage. The turbine component 40 may include a honeycomb structure as shown in Figures 15-17.

[0040] As is illustrated in FIG 4, the tip shroud 26 may include a shroud ridge 42 extending across a portion of the radially outer surface 34 of the shroud 26. The shroud ridge 42 may extend across the pressure side 18 and suction side 20 of the airfoil 12 in a substantially rectangular shape. A plurality of bladelets 44 may connect span-wise along the shroud ridge 42 in a row. The plurality of bladelets 44 produce a plurality of passages 46 with a passage in between each bladelet 44 and along the ends of the shroud ridge 42 for the working fluid. An intershroud bridge 48 may connect each tip shroud 26 between turbine blades 10. The intershroud bridge 48 may extend partially along the ends of the shroud ridge 42. Due to the length of the shroud ridge 42 and the length of the intershroud bridge 48, conventionally half-width passages 50 may be produced along each end of the tip shroud 26 along the extreme ends 52 of the row of plurality of bladelets 44 as is shown in FIG 4. FIG 8 shows a more detailed view of the half-width passages 50 that are produced along the pressure side end 18 and the suction side end 20.

[0041] Embodiments of the present invention provide an inventive technique for accommodating asymmetric flow across the tip shroud 26 with flow guiding features, thus minimizing losses. Embodiments of the present invention provide an inventive technique for efficient turning of the flow field while reducing mixing losses.

[0042] As is shown in FIG 5 and FIG 10, the half- width passages 50 may be covered or blocked. Covering the half-width passages 50 adjacent to at least one of the extreme ends 52 may reduce the separation regions in bladelets 44. The intershroud bridge 48 may be extended so that it is flush with at least one of the bladelets 44 along the shroud ridge 42, thereby covering the half- width passage along that particular end of the airfoil 12. In certain embodiments, the intershroud bridge 48 may cover one end of half- width passages 50. In certain embodiments, the intershroud bridge 48 may cover both ends of half-width passages 50.

[0043] FIG 10 illustrates that in certain embodiments, the intershroud bridge 48 may also cover at least one full passage of the plurality of passages 46. In certain embodiments, the intershroud bridge 48 may cover at least one passage along the suction side 20 of the blade 10. FIG 10 also shows that in certain embodiments, the height (h) of the shroud ridge 42 may also be reduced through any remaining passages 46 in between bladelets 44 so that the total area over the cavity 36 may remain unchanged. A casing 54 may partially cover the cavity 36 within the plurality of passages 46.

[0044] FIG 6 illustrates non-uniform passage widths across the shroud seal 38. The pitch/chord ratio may be non-uniform for the plurality of bladelets 44. An angle of incidence of the plurality of bladelets 44 may be staggered. Varying widths of passages 46 along the shroud seal 38 may allow for the accommodation of asymmetric flow across the tip shroud 26. [0045] The bladelet tips may have a slightly smaller radius as the intershroud bridge 48, so that the intershroud bridge 48 may rub the casing 54 first during transients.

[0046] The scallop region of the tip shroud 26 as is shown in FIG 7 may be reduced. A heavily scalloped tip shroud 26, such as illustrated in FIG 7, may increase parasitic tip leakage and may further distort the streamlines in the outer diameter flow path of the working fluid, increasing aerodynamic losses thereby reducing the stage efficiency. There are advantages to the scallop region, however, so to reduce parasitic tip leakage, additional material may be included around the blade tip 22 along the scallop region to provide for a flow that is more uniform through the bladelet passages 46. The additional material to the scallop region may increase the length upstream of the bladelets for the gas to develop a more uniform and steady flow field. Thus, better inlet conditions to the bladelets 44 may be met by this addition.

[0047] FIG 9 shows another change to the tip shroud 26 that may be introduced. The bladelets 44 are positioned above the shroud ridge 42. Sharp corners on the shroud 26 and the shroud ridge 42 between bladelets 44 may be chamfered to reduce separation. In certain embodiments, the chamfering may be focused on bladelet passages 46, or passages 46 in between the outer two bladelets 44 on each side as is shown in FIG 10.

[0048] FIG 12 and FIG 13 show a view from the leading edge 14 towards the trailing edge 16. Tip leakage may flow over a tip of the seal 38. A filleted tip seal section may provide a redirection of air 64 forward as is shown in FIG 13. The filleted tip seal may reduce the tip leakage flow. The redirection of the flow may create additional blockage that may effectively reduce the tip gap. FIG 14 shows a curved section that redirects air 64 radially inward with radial direction and the rotational direction being highlighted. Tangential bowing of the plurality of bladelets 44 is shown. Another embodiment may include both the scooped tip and the forward facing curved section.

[0049] Changing the airfoil 12 cross-section may reduce efficiency penalties due to low Reynolds number effects by optimizing the airflow across the bladelets 44. These penalties include thickening of boundary layers which increase bladelet profile loss and aerodynamic blockage in bladelet passages.

[0050] Eliminating the half-width bladelet passages 50 along at least one edge may reduce separation regions in the bladelet area. Reducing separation regions may provide a more uniform flow. With a more uniform flow, flow is more likely to remain within the bladelet passages 46, thus providing more work and increasing bladelet efficiency.

[0051] As mentioned above, extending the intershroud bridge 48 to cover, or block, the half- width passages 50 along the extremes of the bladelet 44 rows may reduce tip leakage to a minimum. Separation at the bladelet 44 row inlet may be reduced by chamfering the shroud 26 and shroud ridge 42. In certain embodiments, an additional passage may be blocked off. This additional passage may be on the suction side 20 of the blade 10. The blocking of the additional passage may reduce the flow area for the working fluid. To offset this reduction in flow area, the shroud ridge 42 height may also be reduced through the remaining passages 46. In certain embodiments, the shroud 26 and shroud ridge 42 may be chamfered with the additional blocked passage, and in other embodiment, the shroud 26 and shroud ridge 42 corners are not adjusted.

[0052] 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.