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Title:
GAS TURBINE BLADE TIP SHROUD SEALING AND FLOW GUIDING FEATURES
Document Type and Number:
WIPO Patent Application WO/2017/155497
Kind Code:
A1
Abstract:
A turbine blade (10) includes an airfoil (12) extending span-wise along a radial direction (Ra) relative to a turbine axis having a leading edge (14) and a trailing edge (16) joined by a pressures side (18) and a suction side (20), a tip (22), and a root end (50), and a shroud (26) positioned along the tip (22) of the airfoil (12) extending generally along a circumferential direction (C) relative to a turbine axis. The shroud (26) includes an upstream edge (28) and a downstream edge (30) spaced apart axially, an outer surface (46), and a radially inner shroud surface (52). The shroud (26) further includes a seal (38) extending radially outward from the outer surface (46) of the shroud (26), and an axial extension (36) extending axially upstream from along a side of the seal (38). A method for reducing overtip leakage along the tip shroud comprises a honeycomb structure (42, 44) on a turbine component (40).

Inventors:
LEE, Ching-Pang (12 Camargo Pines Lane, Cincinnati, Ohio, 45243, US)
THAM, Kok-Mun (2592 Ekana Drive, Oviedo, Florida, 32765, US)
Application Number:
US2016/021129
Publication Date:
September 14, 2017
Filing Date:
March 07, 2016
Export Citation:
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Assignee:
SIEMENS AKTIENGESELLSCHAFT (Wittelsbacherplatz 2, München, München, DE)
International Classes:
F01D5/20; F01D5/22; F01D11/12
Foreign References:
EP0957237A21999-11-17
US20110002777A12011-01-06
US3314651A1967-04-18
DE485833C1929-11-08
US20130202439A12013-08-08
US20140147250A12014-05-29
Other References:
None
Attorney, Agent or Firm:
LYNCH, Carly W. (Siemens Corporation- Intellectual Property Dept, 3501 Quadrangle Blvd Ste 230Orlando, Florida, 32817, US)
Download PDF:
Claims:
CLAIMS

What is claimed is:

1. A blade (10) for a turbine engine (32) comprising:

an airfoil (12) extending span-wise along a radial direction (Ra) relative to a turbine axis comprising a leading edge (14) and a trailing edge (16) joined by a pressure side (18) and a suction side (20), a tip end, and a root end (50); and

a shroud (26) positioned along a tip (22) of the airfoil (12) extending generally along a circumferential direction relative to a turbine axis, the shroud (26) comprising:

an upstream edge (28) along the same side as the leading edge of the airfoil, a downstream edge (30) along the same side as the trailing edge of the airfoil (12), an outer surface (46), and a radially inner shroud surface (52);

a seal (38) extending radially outward from the outer surface (46) of the shroud (46);

an axial extension (36) extending axially upstream from a side of the seal (38) and comprising an inner surface (48) facing towards a turbine centerline (11) and an outer surface (56) facing outwardly away from the centerline (11).

2. The blade according to claim 1, wherein the seal (38) comprises a knife edge.

3. The blade according to any of claims 1 or 2, wherein the seal runs a tight gap with a turbine component (40) of the turbine engine comprising a honeycomb structure comprising an axially extended upper portion (44) and a radially inwardly directed extended lower portion (42).

4. The blade according to any of claims 1-3, further comprising a plurality of pumping cutouts (34) on the inner surface (48) of the axial extension (36) angled inwards.

5. The blade according to any of claims 1-3, further comprising a plurality of pumping cutouts (34) along a forward side of the axial extension (36) of the radial tip seal (38) of the blade (10).

6. A method for reducing overtip leakage along a tip shroud (26) with flow guiding features comprising:

providing a blade (10) for a turbine engine (32) comprising:

an airfoil (12) extending span-wise along a radial direction (Ra) relative to a turbine axis comprising a leading edge (14) and a trailing edge (16) joined by a pressure side (18) and a suction side (20), a tip end, and a root end (50); and

a shroud (26) positioned along a tip (22) of the airfoil (12) extending generally along a circumferential direction (C) relative to a turbine axis, the shroud (26) comprising:

an upstream edge (28) along the same side as the leading edge (14) of the airfoil (12), a downstream edge (30) along the same side as the trailing edge (16) of the airfoil (12) spaced apart from each other in an axial direction relative to a turbine axis, an outer surface (46), and a radially inner shroud surface (52);

a seal (38) extending radially outward from the outer surface (46) of the shroud (26);

an axial extension (36) extending axially upstream from the seal (38) and comprising an inner surface (48) facing towards a turbine center line (11) and an outer surface (56) facing outwardly away from the centerline (11); providing a turbine component (40) of the turbine engine comprising a honeycomb structure comprising an axially extended upper portion (44) and a radially inwardly directed extended lower portion (42), wherein in a cold position the radially inwardly directed extended lower portion (42) is closer to a centerline

7. The method according to claim 6, wherein the seal (38) comprises a knife edge and runs a tight gap with the turbine component (40).

8. The method according to any of claims 6-7, wherein the axial extension further comprises a plurality of pumping cutouts (34) on an inner surface of the axial extension (36) angled inwards.

Description:
GAS TURBINE BLADE TIP SHROUD SEALING AND FLOW GUIDING FEATURES

BACKGROUND 1. Field

[0001] 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

[0002] 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.

[0003] 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.

[0004] 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 stage blades. 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.

[0005] 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 restrict 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. Tip leakage is essentially lost opportunity for work extraction. The tip leakage also contributes towards aerodynamic secondary loss.

SUMMARY

[0006] In one aspect of the present invention, a blade for a turbine engine comprises: an airfoil extending span- wise along a radial direction relative to a turbine 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; and a shroud positioned along a tip of the airfoil extending generally along a circumferential direction relative to a turbine 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, an outer surface, and a radially inner shroud surface; a seal extending radially outward from the outer surface of the shroud; an axial extension extending axially upstream from a side of the seal and comprising an inner surface facing towards a turbine centerline and an outer surface facing outwardly away from the centerline. [0007] In another aspect of the present invention, a method for reducing overtip leakage 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 relative to a turbine 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; and a shroud positioned along a tip of the airfoil extending generally along a circumferential direction relative to a turbine 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 spaced apart from each other in an axial direction relative to a turbine axis, an outer surface, and a radially inner shroud surface; a seal extending radially outward from the outer surface of the shroud; an axial extension extending axially upstream from the seal and comprising an inner surface facing towards a turbine centerline and an outer surface facing outwardly away from the centerline; providing a turbine component of the turbine engine comprising a honeycomb structure comprising an axially extended upper portion and a radially inwardly directed extended lower portion, wherein in a cold position the radially inwardly directed extended lower portion is closer to a centerline of the gas turbine engine than the axial extension.

[0008] 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

[0009] 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.

[0010] 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.

[0011] FIG 2 is a detailed perspective view of a turbine airfoil tip shroud and cavity configuration in cold engine conditions of an exemplary embodiment of the present invention.

[0012] FIG 3 is an axial cross sectional view of an axial extension of an exemplary embodiment of the present invention.

[0013] FIG 4 is a detailed perspective view of a turbine airfoil tip shroud and cavity configuration in hot engine conditions of an exemplary embodiment of the present invention.

DETAILED DESCRIPTION [0014] 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.

[0015] Broadly, an embodiment of the present invention provides a turbine blade that includes an airfoil extending span-wise along a radial direction relative to a turbine axis having a leading edge and a trailing edge joined by a pressures side and a suction side, a tip, and a root end, and a shroud positioned along the tip of the airfoil extending generally along a circumferential direction relative to the turbine axis. The shroud includes an upstream edge and a downstream edge spaced apart axially, an outer surface, and a radially inner shroud surface. The shroud further includes a seal extending radially outward from the outer surface of the shroud and an axial extension extending axially upstream from along a side of the seal.

[0016] 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.

[0017] Any tip leakage flow 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. [0018] A reduction in gas leakage across blade tip and increases 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.

[0019] 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 the working fluid. The blades 10 are rotatable about a rotation axis along a centerline 11 of the gas turbine engine 32. The blade rotation Ro is shown in Figure 3. Each blade 10 is formed from an airfoil 12 extending span-wise in a radial direction Ra relative to the turbine axis in the turbine engine 32 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 a root end 50 of 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.

[0020] 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. A radially inner surface 52 of the shroud 26 adjoins the tip 22 of the airfoil 12. The shroud 26 includes an outer surface 46 opposite of the radially inner surface 52. The radially inner surface 52 and the outer surface 46 are connected by the upstream edge 28 and the downstream edge 30. [0021] FIG 2 and FIG 4 show the area around the tip shroud 26 and a cavity 54 configuration in a detailed view. The shroud 26 may extend along a circumferential direction relative to the turbine axis. A seal 38 may be provided on the shroud 26, extending radially outward from the outer surface 46 of the shroud 26 away from the centerline 11. 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 32 reducing overtip leakage. The turbine component 40 may include a stepped honeycomb structure such as shown in Figures 1, 2, and 4. The stepped honeycomb includes an axially extended upper portion 44 and a radially inwardly directed extended lower portion 42. The turbine component 40, referred to as the stepped honeycomb, may be on a stator, such as a ring segment. When referring to an outward direction, this refers to a direction away from the centerline 11 of the gas turbine engine 32. When referring to an inward direction, this refers to a direction towards the centerline 11 of the gas turbine engine 32.

[0022] Embodiments of the present invention provide an inventive technique for accommodating 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.

[0023] Additionally, the shroud 26 may include an axial extension 36. The axial extension 36 may extend in the axial direction (A) upstream from the seal 38. The axial extension 36 may include an inner surface 48 facing towards the centerline 11 and an outer surface 56 facing outwardly away from the centerline 11. The axial extension 36 provides an additional outcropping from the shroud 26 that interacts with the radially inwardly directed extended lower portion 42 of the stepped honeycomb. The stepped honeycomb may initially extend further towards the centerline 11 than where the axial extension 36 positioned.

[0024] In certain embodiments, a plurality of pumping cutouts 34 may be cutout along the inner surface 48 of the axial extension 36 of the shroud 26 as is shown in Figure 3. The plurality of pumping cutouts 34 along the inner surface 48 of the axial extension 36 may provide a redirection of gas downward and radially inward. This redirection may reduce the gas flow up and over for a tip leakage flow across and past the radial seal 38 by reducing the total amount of air flow directed towards the cavity 54. In certain embodiments, the plurality of pumping cutouts 34 may be cast as an integral feature of the blade 10. [0025] The plurality of pumping cutouts 34 may have one of several different shapes in order to fit an application. The plurality of pumping cutouts 34 may have a predominately triangular shape, deeper as the cutouts 34 move axially downstream, deeper as the cutouts 34 move axially upstream, have both straight edges and curves, tubular, or the like. The size and shape of each cutout 34 may be determined by mechanical and aerodynamic requirements such as the airfoil radial growth and untwist at operating conditions. The plurality of pumping cutouts 34 can be any shape that may be required for the axial extension 36 geometry and damping characteristics as long as the plurality of pumping cutouts 34 allow for a forcing of air radially inward. [0026] The seal 38 and axial extension 36 may work in conjunction with the stepped honeycomb. 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 Figure 2. The seal 38 and axial extension 36 are positioned below or inward from the stepped honeycomb. There is a cavity 54 that is created which is the space between the seal 38 and the axially extended upper portion 44 of the stepped honeycomb. The cavity 54 spacing should be kept to a minimum in in order to reduce gas leakage across the seal 38 in a circumferential direction C. In this cold position, the stepped honeycomb is uncut.

[0027] 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 4, the stepped honeycomb is cut by both the axial extension 36 and the seal 38 along the tip shroud 26. The resultant stepped honeycomb gouges are shown in Figure 4. The gas turbine engine 32 in a hot running position with the seal 38 relative to the honeycomb is also shown in Figure 4. Two sealing points may be created. The seal 38 may provide a sealing point along the axially extended upper portion 44 of the honeycomb structure. The axial extension 36 may provide another sealing point along the radially inwardly directed extended lower portion 42 of the honeycomb structure. The seal 38 and the axial extension 36 may exploit the relative position of the rotor with respect to the stator during steady- state conditions. The rotor tends to move towards the stator, due to either the use of hydraulic clearance optimization or a rotor air cooler. The rotor air cooler has the effect of placing the rotor forward relative to the stator, as the axial distribution of the rotor temperature is cooler than the casing parts. The plurality of pumping cutouts 34 on the inner surface 48 of the axial extension 36 may further reduce tip leakage by deterring radial outflow of leakage air at the axial extension 36/stepped honeycomb interface.

[0028] The gas turbine engine blade may have a higher turbine aerodynamic efficiency and improved sealing with the addition of the axial extension 36. Adding the stepped honeycomb may further improve the sealing around the tip of the airfoil 12.

[0029] 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.