STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH

[0001] Development of this invention was supported in part by the United States Department of Energy, Contract No. DE-FC26-05NT42644. Accordingly, the United States Government may have certain rights in this invention.

BACKGROUND

1. Field

[0002] The present invention relates to turbine engines, and more specifically to a honeycomb abradable feature for gas turbine tip shrouded blades. 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] Better engine performance requires higher component efficiencies such as in the compressor and turbine blading. 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 are the last several stage blades and between rotating blades (the interstage area). Traditionally the last stage blades have 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] In current assemblies, the rotating blade tip shroud and cavity configurations in large industrial gas turbines are regions of low performance. Another area of low performance is interstage areas where a rotor has rotating knife- edge seals rubbing against a stationary shroud under a vane. There are several drivers of aerodynamic loss in any turbine-shroud cavity configuration, which lowers the gas turbine’s efficiency. One driver is the flow over the rotating tip of the seal. Tip seals are generally designed to restrict the flow and consequently lead to high flow velocities in the 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. Any leakage is essentially lost opportunity for work extraction. The leakage also contributes towards aerodynamic secondary loss.

SUMMARY

[0007] In one aspect of the present invention, a turbine engine comprises: at least one blade comprising an airfoil extending span-wise along a radial direction relative to a turbine rotation axis comprising a leading edge and a trailing edge joined by a pressure side and a suction side, a tip end, and a root end opposite of the tip end radially closer to the turbine rotation axis than the tip end; at least one vane comprising an airfoil extending span-wise along a radial direction relative to a turbine rotation axis comprising a leading edge and a trailing edge joined by a pressure side and a suction side, a tip end, and a root end opposite of the tip end radially closer to the turbine rotation axis than the tip end; at least one stationary shroud positioned radially inward from the vane extending generally along a circumferential direction relative to a turbine rotation axis; a seal extension extending radially outward towards the shroud from a rotor; at least one shroud positioned along the tip end of the airfoil of the at least one blade extending generally along a circumferential direction relative to a turbine rotation axis, the shroud comprising: an upstream edge along the same side as the leading edge of the airfoil, a downstream edge along the same side as the trailing edge of the airfoil, a radially outer surface, and a radially inner surface; a seal extension extending radially outward from the radially outer surface of the shroud, wherein the seal extension runs in a circumferential direction along the radially outer surface of the shroud; and at least one honeycomb abradable structure attached to at least one stationary component of the turbine engine, wherein the honeycomb abradable structure is canted and facing at least one seal extension.

[0008] In another aspect of the present invention, a method for reducing leakage over a seal extension with flow guiding features comprises: installing a plurality of blades and vanes for a turbine engine each comprising: an airfoil extending span-wise along a radial direction relative to a turbine rotation axis comprising a leading edge and a trailing edge joined by a pressure side and a suction side, a tip end, and a root end opposite of the tip end radially closer to the turbine rotation axis than the tip end; wherein the turbine engine additionally comprises: at least one stationary shroud positioned radially inward from the vane extending generally along a circumferential direction relative to a turbine rotation axis; a seal extension extending radially outward towards the shroud from a rotor; at least one shroud positioned along the tip end of the airfoil of the at least one blade extending generally along a circumferential direction relative to a turbine rotation axis, the shroud comprising: an upstream edge along the same side as the leading edge of the airfoil, a downstream edge along the same side as the trailing edge of the airfoil, a radially outer surface, and a radially inner surface; a seal extension extending radially outward from the radially outer surface of the shroud, wherein the seal extension runs in a circumferential direction along the radially outer surface of the shroud; and installing a honeycomb abradable structure of the turbine engine comprising a canted form extending radially inward from a stationary component of the turbine engine, wherein the honeycomb abradable structure is canted to reduce a clearance volume between the honeycomb abradable structure and the seal extension.

[0009] 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 [0010] 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.

[0011] FIG 1 is an axial cross-sectional view of a gas turbine engine with a row of shrouded turbine blades wherein embodiments of the present invention may be incorporated.

[0012] FIG 2 is a perspective view of a turbine blade as part of a turbine engine.

[0013] FIG 3 is a partial perspective view of two blade tips.

[0014] FIG 4 is a perspective top view of a honeycomb abradable structure as part of a turbine engine of the prior art. [0015] FIG 5 is a partial perspective view of a honeycomb abradable feature of an exemplary embodiment of the present invention for an interstage seal area.

[0016] FIG 6 is a partial front view of a honeycomb abradable feature of an exemplary embodiment of the present invention for an interstage seal area.

[0017] FIG 7 is a chart of pressure ratio versus flow function comparison between a baseline prior art and an exemplary embodiment of the present invention.

[0018] FIG 8 is a chart of pressure ratio versus flow function comparison between a baseline prior art and an exemplary embodiment of the present invention.

DETAILED DESCRIPTION

[0019] 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. [0020] Broadly, an embodiment of the present invention provides a turbine engine that includes at least one blade that includes an airfoil extending span-wise along a radial direction. The turbine engine includes a shroud positioned along a tip end of the airfoil extending along a circumferential direction relative to the turbine rotation axis. The shroud includes an upstream edge and a downstream edge spaced apart axially, a radially outer surface, and a radially inner surface. A seal extension extends radially outward from the radially outer surface of the shroud and runs in a circumferential direction along the radially outer surface of the shroud. A canted honeycomb abradable structure attaches to a stationary component of the turbine engine.

[0021] 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.

[0022] Any leakage flow is lost work extraction, thus lowering the turbine efficiency. One area of concern is the flow over the outer radial tip of the rotating blade. Tip shrouds can be used. Tip shrouds are placed along the tip of the blade with seals to help reduce tip leakage. When there are rotating parts such as a blade or disk that are adjacent to a stationary part such as a casing, there is a need for a material that the rotating part can rub against and wear away. This material allows for the protection of the main structures, both rotating and static, from damage and loss of sealing.

[0023] High temperature regions of the gas turbine can include abradable material. An example is FIG. 4 of a metallic honeycomb. The metallic honeycomb can be made by forming thin strips of sheet metal to a half-hexagon shape, then spot welding the formed strips together to form a straight six-sided cell as shown. The cellular structure can then be brazed to a backing plate. The resultant material can be stiff, light, easily rubbed into, and cut by the rotating part. The rotating part can have a labyrinth or knife edge seal. The interaction between the individual cells plus the knife edge provide a loss mechanism to the flow by progressive expansions and contractions that reduce pressure within the seal cavity that minimizes the leakage. The seal is still subject to leakage that increases at higher pressure ratios. The flow within the seal can be difficult to predict, with jets, and recirculation that can reduce the sealing effectiveness. One or more serrations, i.e. labyrinth or knife edge sealing, rub against the honeycomb.

[0024] A labyrinth type seal can be beneficial by absorbing energy in the gas that goes towards forcing a leak. If there is more pressure on one side and relatively little pressure on the other side, the driving pressure will cause a leak. Any labyrinth seal is a loss producing mechanism. The labyrinth seal works so that there is so much of a loss that virtually no flow gets through.

[0025] Tip leakage control and sealing efficiency is desirable. Embodiments of the present invention provide a canted honeycomb abradable structure for tip shrouded blades that may allow for the reduction in losses. Canting providing angled oblique lines that cut off a corner. The angle being less than a right angle.

[0026] Referring to FIG 1, a portion of a turbine section of a gas turbine engine 32 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 working fluid. The flow of the working fluid Wf is in the direction as shown in Fig. 1. The blades 10 are rotatable about a turbine rotation axis along a centerline 11 of the gas turbine engine 32. Each blade 10 is formed from an airfoil 12 extending span-wise in a radial direction Ra relative to the turbine rotation axis along the centerline 11 in the turbine engine 32 from a rotor disc 42. The airfoil 12 includes a leading edge 14 and a trailing edge 16 joined by a pressure side 18 and a suction side 20 on a side opposite to the pressure side 18, a tip end 22 at a radially outer end of the airfoil 12 away from the centerline 11, a platform 24 coupled to a root end 34 of the airfoil 12 at a radially inner end of the airfoil 12 towards the centerline 11, for supporting the airfoil 12 and for coupling the airfoil 12 to the rotor disc. The blade 10 further includes a shroud 26, referred to as a tip shroud, coupled to the tip end 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.

[0027] The shroud 26 includes an upstream edge 28 along the same side as the leading edge 14 of the airfoil 12 and a downstream edge 30 along the same side as the trailing edge 16 of the airfoil 12. The term upstream and downstream are in relation to the flow of the working fluid Wf. A radially inner surface 46 shown in FIG. 3 of the shroud 26 adjoins the tip end 22 of the airfoil 12. The shroud 26 includes a radially outer surface 36 shown in FIG. 3 opposite of the radially inner surface 46. The radially inner surface 46 and the radially outer surface 36 of the shroud 26 are connected by the upstream edge 28 and the downstream edge 30. The shroud 26 can follow a tapered angle along the tip end 22 of the airfoil 12 in some embodiments. FIGs 1 through 3 further show a seal extension 38 that extends radially outward from the tip shroud 26. Shown is a knife edge seal extension 38 and is positioned along the radially outer surface of the tip shroud 26 with an end along a circumferential leading end and a circumferential trailing end adjacent to the adjacent knife edge seal extension 38. The seal extension 38 can be knife edge or other shape seal. Another knife edge seal extension 38 is shown on the rotor 42 below a stationary shroud 50 under a vane 48. The knife edge seal extension 38 in this location is also rotating so that it rubs against the stationary shroud 50 under the vane 48. FIG. 1 shows a honeycomb abradable structure 40 radially outward from the knife edge seal extension 38 in both locations. The seal extension 38 can be stepped to provide more of a labyrinth seal as is shown in the knife edge seal extension 38 along the rotor. Also pointed out in FIG. 1 is the disc cavity 44 between the vane 48 and blade 10. There can be at least one honeycomb abradable structure 40 attached to the stationary shroud 50 radially inward from the vane 48 facing the seal extension 38 along the rotor 42 and/or above the tip shroud 26 of the blade 10.

[0028] Like the blades 10, the vanes 48 also include an airfoil 12 extending span- wise along a radial direction Ra relative to a turbine rotation axis 11 comprising a leading edge 14 and a trailing edge 16 joined by a pressure side 18 and a suction side 20, a tip end 22, and a root end 34 opposite of the tip end 22 radially closer to the turbine rotation axis 11 than the tip end 22. As mentioned above, the stationary shroud 50 is radially inward from the vane 48 extending generally along the circumferential direction C relative to a turbine rotation axis.

[0029] FIGS 2 and 3 show blades with a shroud 26. Along the outer surface 36 of the tip shroud 26 is the knife edge seal extension 38. As is shown, the knife edge seal extension 38 can run the length of the tip shroud 26 along the circumferential direction. The flow of the working fluid Wf is shown to demonstrate that the knife edge seal extension 38 in relation to the flow of the working fluid Wf.

[0030] Honeycomb structures have been made with the cells in a straight radial orientation in the past. Due to spot welding and the like, this has been the way honeycombs are made.

[0031] The flow that wants to jet through the knife edge seals is not a straight axial flow since the rotor is rotating. A shearing force is produced that causes the leakage flow to swirl slightly. The flow follows a three-dimensional pattern. The honeycomb abradable structure 40 is reoriented to maximize the sealing. If the honeycomb is angled against the direction of flow it can catch the flow the same way as a canted knife edge seal would. The problem had been viewed in two dimensions, but the problem is actually in a three-dimensional spacing. By canting the honeycomb abradable structure 40, maximizing expansion and contraction losses and improving sealing efficiency can be accomplished. The charts in FIG. 7 and FIG. 8 show increases reduction of leakage flow across the blade.

[0032] FIG. 7 shows the benefit of multiple features together. The line that is shown mostly below the second line includes adjustments of canting the honeycomb as well as stepping the seal. The standard design with no steps or canting is the other line shown using the same radius. The pressure ratio is the pressure of the high side to the low side that is a measure of the energy. FIG. 8 shows performance improvement of a canted honeycomb abradable structure 40 over a conventional honeycomb. FIG. 8 shows as the pressure ratio goes up there is a constant benefit, but as the ratio goes down the benefit goes away for the conventional honeycomb. At this point there is more driving force to force the leak. While the canted honeycomb has a relatively constant performance. As the pressure ratio goes up the baseline design is less and less effective as a seal because of more energy. It becomes overwhelmed. As higher pressure ratios come in to play, an embodiment of the honeycomb abradable structure 40 is more effective. Since the canted honeycomb is more efficient at the higher pressure ratios, seals with fewer teeth can be used. Achieving the same sealing function with three teeth instead of four for example can make the seal smaller and more compact and cost less money. As the pressure ratio goes up, typically, a longer seal would be required. With the honeycomb abradable structure 40 a longer seal is now not required for higher pressure ratios. Further, flexibility is available now in regards to space for example. A seal with four teeth may be desirable, but there may be only space for three. In this case, the honeycomb abradable structure 40 allows for the sealing needed.

[0033] Conventional technologies require capital intense manufacturing to make a conventional honeycomb. As mentioned above, made by forming thin strips of sheet metal in straight radial orientation to a half-hexagon shape, then spot welding the formed strips together to form a six-sided cell as shown. Brazing the cellular structure to a backing plate. Variability was limited based on manufacturing restrictions. The honeycomb abradable structure 40 discussed here within can be manufactured by directly depositing metal onto a backing plate using additive manufacturing. Due to the direct depositing, any orientation can be used with this method of manufacturing. The honeycomb abradable structure 40 can be canted versus the conventional straight cell structure. The honeycomb abradable structure 40 can be canted at an angle. The angle produced can drive the flow away from the further side of the blade, thereby increasing the pressure drop, improving efficiency. The honeycomb abradable structure 40 can be canted in a compound angle that has circumferential and axial components.

[0034] FIG. 1 shows a placement of a shrouded blade in a traditional positioning. A honeycomb abradable structure 40 is positioned between the shrouded blade and the casing of the turbine engine 32. The honeycomb abradable structure 40 extends in a radially inward direction towards the airfoil 12. A honeycomb abradable structure 40 is also shown positioned between the knife edge seal extension 38 of the rotor 42 and the shroud 50 of the vane 48. FIG 1 shows a cold state of the turbine engine 32. As the turbine heats up, the edge cuts into the honeycomb structure 40. [0035] FIG. 5 and FIG. 6 illustrate an embodiment of the present invention. The honeycomb abradable structure 40 is positioned as a canted component. The canting of the honeycomb abradable structure 40 in these FIGS are opposite the angle of the knife edge seal extension 38. The canted honeycomb abradable structure 40 in certain embodiments is also canted in the circumferential direction C. The canting in the circumferential direction reduces the clearance impact to aerodynamic performance. The canted honeycomb abradable structure 40 allows for a tight fit between the tip shroud 26 and the honeycomb abradable structure 40 along a blade chord line. The canted honeycomb abradable structure 40 provides flow guidance through the placement of the canted honeycomb abradable structure 40 which will be described below. The blade chord line defined as the path of an imaginary straight line joining the leading 14 and trailing edges 16. Less space is provided for leakage to cross over when the honeycomb abradable structure 40 is canted since the honeycomb abradable structure 40 is oriented for the flow through the spacing coming over the tip of the blade. The flow, moving in three dimensions, is slowed down by the tortuous path from the honeycomb abradable structure 40.

[0036] Embodiments of the present invention provide an inventive technique for accommodating flow across the tip shroud 26 with flow guiding features of a canted honeycomb abradable structure 40, thus minimizing losses. The honeycomb abradable structure 40 can provide a seal with fewer teeth than traditionally required. Smaller and more compact honeycomb abradable structures 40 can be produced at lower costs in order to provide the same or better sealing of the area above the tip of the blade.

[0037] When the gas turbine engine 32 is in a cold position, such as when the gas turbine engine 32 is not in operation, the positioning of the components may be as shown in FIG. 1. The tip shroud 26 is positioned below or radially inward from the canted honeycomb abradable structure 40.

[0038] Once the gas turbine engine 32 is running and the blade is in a hot position, from hot and warm restarts, such as is shown in Figure 8, the canted honeycomb abradable structure 40 is cut by the knife edge seal extension 38 along the tip shroud 26 or some labyrinth seal. Having the canting of the honeycomb abradable structure 40 generally opposite the angle of the tip shroud 26 allows for a minimal amount of working fluid to flow across and over the tip shroud 26 reducing the cavity or clearance between the honeycomb abradable structure 40 and the tip shroud 26 along the blade 10. [0039] The gas turbine engine blade may have a higher turbine aerodynamic efficiency and improved sealing with the addition of the canted honeycomb abradable structure 40. Adding the canted honeycomb abradable structure 40 with stepped seal further improves the sealing around the tip end 22 of the airfoil 12.

[0040] 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.