WO2017039568A1

WO2017039568A1 - Turbine airfoil cooling channel with fenced pedestals

Info

Publication number: WO2017039568A1
Application number: PCT/US2015/047323
Authority: WO
Inventors: George Liang; Nan Jiang; Jan H. Marsh; Ching-Pang Lee; Jae Y. Um
Original assignee: Siemens Aktiengesellschaft; Siemens Energy, Inc.
Priority date: 2015-08-28
Filing date: 2015-08-28
Publication date: 2017-03-09

Abstract

An air cooled turbine airfoil and method of cooling includes an outer wall (18) including pressure side wall and a suction side wall extend between a leading edge (12) and a trailing edge (14). A multiple pass serpentine flow cooling circuit extends along the turbine airfoil aft from the leading edge (12) region to the trailing edge (14) region. A plurality of pedestals (26) extends across at least one of the channels of the serpentine flow cooling circuit and is spread apart in a chordal direction to define radial passages for aft flowing cooling air (20). Discrete trip strips (30) are incorporated on the plurality of pedestals (26) forming a plurality of fenced pedestals (26). At least a portion of the quantity of plurality of pedestals (26) is tapered in length from a large diameter end (38) to a small diameter end (40). The small diameter end (40) of the plurality of pedestals (26) is adjacent to the outer wall (18) of the airfoil.

Description

TURBINE AIRFOIL COOLING CHANNEL WITH FENCED PEDESTALS

BACKGROUND 1. Field

[0001] The present invention relates to gas turbine engines, and more specifically to a turbine airfoil cooling channel with fenced pedestals.

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 to the material properties and cooling capabilities of the turbine parts. This is especially important for first stage turbine vanes and blades since these airfoils are exposed to the hottest gas flow in the system. [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] Since the turbine blades are exposed to the hot gas flow discharged from combustors within the combustion system, cooling methods are used to obtain a useful design life cycle for the turbine blade. Blade cooling is accomplished by extracting a portion of the cooler compressed air from the compressor and directing it to the turbine section, thereby bypassing the combustors. After introduction into the turbine section, this cooling air flows through passages or channels formed in the airfoil portions of the blades. [0005] In order to allow for higher temperatures, turbine blade designers have proposed several complex internal blade cooling circuits to maximize the blade cooling through the use of convection cooling, impingement cooling and film cooling of the blades. Figure 3 shows a prior art super cooled turbine blade design.

[0006] FIG. 4 shows a basic flow phenomenon over a pedestal and the heat transfer performance around a single pedestal. The heat transfer performance for a current pedestal can be estimated with the use of a cylinder in a cross flow model. Cooling flow encountering the front portion of the cylinder yields a high heat transfer coefficient, especially at the stagnation point. The heat transfer performance is then slowly decreased along the pedestal circumference as is shown in Figure 5. At the back end of the pedestal a pair of vortices is generated reducing the speed and thus reducing the heat transfer performance. Higher temperatures are then found along the outer wall of the airfoil versus along an inner partition wall.

[0007] Turbine airfoils (which include blades and vanes) are typically cast as a single piece with the cooling passages cast within the airfoil. Ceramic cores having the cooling passage shape are used to form the airfoil. One problem with the investment casting process that is used to produce a turbine airfoil is that the cooling passages within the airfoil have pin fins that are formed parallel to each other within the common passage or passages formed from a single ceramic core. Because of the die pulling direction in the die that is used to cast the ceramic core, the pin fins are limited to being in the pulling direction of the mold and thus are all parallel to each other. [0008] In the investment casing process, there are minimum wall thicknesses that can be cast because of the viscosity of the molten metal and its capacity to flow through the mold and around the ceramic cores through small holes or spaced. Also, with investment casting only a single metal or alloy can be poured into the mold. Thus, producing a single metallic piece of composite metal materials is not possible with this process.

[0009] The prior art turbine blades produced using the investment casting process also have the blade root is cast without the fir tree configuration for mounting within the slots of the rotor disk. In this process, the blade is cast first and then the fir tree configuration is machined into the root portion. This adds further expense and complexity to the production of a turbine rotor blade. SUMMARY

[0010] In one aspect of the present invention, an air cooled turbine airfoil comprising: a leading edge and a trailing edge; an outer wall comprising a pressure side wall and a suction side wall extending between the leading edge and the trailing edge; a multiple pass serpentine flow cooling circuit extending along the turbine airfoil from the leading edge region to the trailing edge region; and a plurality of pedestals extending across at least one of the channels of the serpentine flow cooling circuit, the plurality of pedestals spread apart in a chordal direction to define radial passages for aft flowing cooling air, wherein at least a portion of the quantity of plurality of pedestals are tapered in length from a large diameter end to a small diameter end, wherein the small diameter end of the plurality of pedestals is adjacent to the outer wall of the turbine airfoil.

[0011] In another aspect of the present invention, a method for increasing a heat transfer coefficient through a cooling circuit of a turbine blade, comprises: providing a turbine airfoil comprising: a leading edge and a trailing edge; an outer wall comprising a pressure side wall and a suction side wall extending between the leading and trailing edges; a multiple pass serpentine flow cooling circuit extending along the turbine airfoil from the leading edge region to the trailing edge region; and a plurality of pedestals extending across at least one of the channels of the serpentine flow cooling circuit, wherein at least a portion of the plurality of pedestals are tapered in length from a large diameter end to a small diameter end, wherein the small diameter end of the plurality of pedestals is adjacent to the outer wall of the turbine airfoil; and sending cooling air through the multiple pass serpentine flow cooling circuit, passing the cooling air around the plurality of pedestals, wherein the tapered shape of the plurality of pedestals forces the cooling air towards the small diameter end with an increased heat transfer coefficient.

[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 super cooled turbine blade.

[0015] FIG 2 is a perspective view of a turbine blade with an airfoil exterior wall removed.

[0016] FIG 3 is a cross sectional top view of a prior art turbine blade internal cooling circuits.

[0017] FIG 4 is a front view of a single pedestal of the prior art.

[0018] FIG 5 is a graph showing heat transfer coefficient performance of a prior art pedestal.

[0019] FIG 6 is a serpentine flow path of a cooling flow circuit.

[0020] FIG 7 is a cross sectional top view of an exemplary embodiment of the present invention along 7-7 in Fig 1.

[0021] FIG 8 is a detailed front view of a tapered pedestal of an exemplary embodiment of the present invention.

[0022] FIG 9 is a perspective view of a tapered pedestal of an exemplary embodiment of the present invention.

[0023] FIG 10 is a detailed front view of a plurality of pedestals of an exemplary embodiment of the present invention.

[0024] FIG 1 1 is a detailed front view of a plurality of pedestals of an alternate embodiment of the present invention.

[0025] FIG 12 is a side view of a tapered pedestal of an exemplary embodiment of the present invention. [0026] FIG 13 is a pressure side view of ceramic core of an exemplary embodiment of the present invention.

[0027] FIG 14 is a suction side view of ceramic core of an exemplary embodiment of the present invention. DETAILED DESCRIPTION

[0028] Broadly, an embodiment of the present invention provides an air cooled turbine airfoil and method of cooling includes an outer wall including pressure side wall and a suction side wall extend between a leading edge and a trailing edge. A multiple pass serpentine flow cooling circuit extends along the turbine airfoil aft from the leading edge region to the trailing edge region. A plurality of pedestals extends across at least one of the channels of the serpentine flow cooling circuit and is spread apart in a chordal direction to define radial passages for aft flowing cooling air. At least a portion of the quantity of plurality of pedestals is tapered in length from a large diameter end to a small diameter end. The small diameter end of the plurality of pedestals is adjacent to the outer wall of the airfoil.

[0029] A blade of a gas turbine receives high temperature gases from a combustion system in order to produce mechanical work of a shaft rotation. Due to the high temperature gases, a cooling system may be provided to reduce the temperature levels throughout the blade. [0030] A gas turbine cooling system may perform at least two basic functions. The first function may be to provide direct cooling of components exposed to gas path temperature that is higher than material temperature limits. The second function may be that of turbine environmental control. Air at correct pressure and temperature may be provided at various critical points to ensure that design environment is maintained throughout the turbine.

[0031] In certain embodiments, air for cooling the rotor and rotating blades may be extracted from the axial compressor discharge at a combustor shell. The compressor discharge air may pass through an air-to-air cooler and may be filtered for rotor cooling. Direct cooling may occur at the turbine spindle blade root end along one or more stages. The turbine stationary vanes may be cooled by both internal bypassing and external bleeding lines.

[0032] An effective step that can be taken to increase the power output and improve the efficiency of a gas turbine engine may be to increase the temperature at which heat is added to the system, that is, to raise the turbine inlet temperature of the combustion gases directed to the turbine. Increases in efficient turbines have led to an increase in the temperature that must be withstood by the turbine blades and rotor. The result is that to use the highest desirable temperatures, some form of forced cooling may be desirable. This cooling may be in the form of air bled from the compressor at various stages, and ducted to critical elements in the turbine. Although emphasis is placed on cooling the initial stages of vanes and blades, air may be also directed to other vanes, blades, ring segments, and discs.

[0033] Better cooling along the outer wall of a blade through additional turbulence and higher heat transfer coefficients is desirable. Embodiments of the present invention provide a blade that may allow for the reduction in temperature along the outer wall of the blade without the use of additional cooling air.

[0034] As is illustrated in Figures 1, 2, and 6 through 14, a turbine rotor blade 10 may include a leading edge 12 and a trailing edge 14. Embodiments of the present invention may also be adapted for use in a turbine vane, both of which are considered to include turbine airfoils. An airfoil is the portion of the blade 10 or vane that reacts with the hot gas flow and has an airfoil cross sectional shape with the leading edge 12 and the trailing edge 14, an outer wall 18 including a pressure side (PS) and a suction side (SS) forming the airfoil shape. The turbine rotor blade 10 may include at least one cooling circuit. The at least one cooling circuit may be a serpentine flow path cooling circuit. The at least one cooling circuit helps move flow of air 20 from the leading edge 12 to the trailing edge 14 in order to help reduce the blade temperature throughout the blade 10. The blade 10 may include an inner partition wall 16. The inner partition wall 16 may extend chord-wise substantially running the length in the airfoil. Figures 1 and 2 show an embodiment of the turbine blade 10 with and without the external airfoil wall shown. [0035] The at least one cooling circuit may include multiple pass cooling channels that are connected through substantially 180-degree turns along a tip end 22 and a root end 24 of the blade 10 that change the direction of the multiple pass cooling channels as the air flow 20 moves aft along the airfoil as shown in Figure 6. [0036] A plurality of pedestals 26, also known as pin fins, may be connected to the airfoil and spread out along the chordal direction (C). The plurality of pedestals 26 may be connected to either the pressure side or the suction side wall while in the mid- chord channels. Along a trailing edge radial out flow cooling channel, the plurality of pedestals may be connected onto both the pressure side and the suction side walls as is shown in Figure 7. At least a portion of the plurality of pedestals 26 may have a tapered shape having a large diameter 38 at one end and a small diameter 40 at another end. In some instances the at least a portion of the plurality of pedestals 26 have a conical shape. The large diameter end 38 may be adjacent to the cold inner partition wall 16, while the small diameter end 40 may be adjacent to the hot outer wall 18. The hot and cold designations are in relation to each other. The tapered shape to the plurality of pedestals 26 may allow the cooling flow to be forced towards the hot outer wall 18, thereby providing more cooling to that area. The plurality of pedestals 26 may define radial passages for the air flow 20 to move through as the air flow 20 moves aft along the airfoil. [0037] Cooling flow of air 20 through the cooling circuit with plurality of pedestals 26 may not become fully turbulent until after having passed through the first few rows of pedestals 26. Initially, pedestal flow may be described by a turbulent flat plate model. As the cooling flow passes successively through the minimum flow cross-sectional area formed by the pedestals 26 it becomes increasingly turbulent and the benefit of pedestal cooling is finally realized.

[0038] In certain embodiments, discrete trip strips 30 may be constructed onto a plurality of pedestals 26. The discrete trip strips 30 may partially obstruct the cooling channels or trip a boundary layer as cooling flow entering the plurality of pedestals 26. The cooling flow may become highly turbulent upon entering a first pedestal row and effective cooling may begin immediately. The cooling performance may be enhanced all around the plurality of pedestals 26 regardless of their flow angle relative to a stagnation point 28. Figure 8 shows an exemplary embodiment of the pedestal with such discrete trip strips 30. Figure 9 shows a view of this design from a perspective view.

[0039] In certain embodiments, a second set of discrete trip strips 32 may be incorporated onto an airfoil inner channel wall 36 from pedestal to pedestal in a staggered array as well as the discrete trip strips 30 on the pedestal surface as is shown in Figure 7. The second set of discrete trip strips 32 may be incorporated on the pedestals 26 along with the discrete trip strips 30, helping to form fenced pedestals 26.

[0040] Figures 10, 1 1 and 12 show embodiments of a flow pattern for cooling air 20 around and past the plurality of pedestals 26 and a view of one of the plurality of pedestals 26 with the discrete trip strips 30 along the sides of the pedestal. Space may be available between the discrete trip strips 30 along the plurality of pedestals 26. As the cooling air 20 flow passes around the plurality of pedestals 26, the boundary layer may be tripped by the second set of discrete trip strips 32. [0041] The discrete trip strips 30 along the sides of the pedestals may be of various combinations and patterns, with Figures 10 and 1 1 showing two of many

embodiments. Figure 10 shows the plurality of pedestals in an equilateral formation. In one embodiment, six discrete trip strips may be spread around the pedestal at approximately 60 degrees in between each discrete trip strip 30. Figure 11 shows a square pattern as an example. In one embodiment, the square pattern of discrete trip strips 30 may have four discrete trip strips spread around the pedestal at

approximately 90 degrees from each discrete trip strip.

[0042] As can be seen in Figures 8 through 12, the discrete trip strips 30 may extend out from the plurality of pedestals 26 in a fenced like format. The fenced plurality of pedestals 26 may partially obstruct the cooling channels and help create additional turbulence as the cooling air flows pass the plurality of pedestals 26.

[0043] Figure 7 shows a cross section view of the turbine airfoil. This view shows the second set of discrete trip strips 32 along the inner channel wall, as well as the inner partition wall 16 and the tapered shaped of the pedestals 26 in between the inner partition wall 16 and the outer wall 18. In certain embodiments, additional cooling holes 34 may be added to the turbine airfoil for the cooling of the airfoil leading edge 12 and the trailing edge 14.

[0044] The turbine blade 10 may be manufactured by direct metal laser sintering (DMLS) technique, 3-D printing a metal blade 10 directly, 3-D printing an integral ceramic core for a lost wax casting method, using a disposable core die to make an integral ceramic core for lost wax casting, assembling multiple ceramic core pieces together to form an integral ceramic core for lost wax casting, or the like.

[0045] The manufacturing process can be used to produce very fine details within metallic structures that cannot be cast using a present day investment casting process. The entire airfoil may be printed as a single piece or printed as a ceramic core. The airfoil and plurality of pedestals 26 may be printed using a high temperature resistant material. The plurality of pedestals 26 may be of a different material than the airfoil. Figures 13 and 14 show embodiments of the ceramic casting core as described. [0046] These different types of manufacturing processes allow for the shaping and direction of the pedestals 26, as well as allowing the discrete trip strips 30 to be added directly to the plurality of pedestals 26. With the use of these manufacturing techniques, all the pedestals 26 need not be oriented in the same direction. The pedestals 26 may be tapered from a small diameter at the hot outer wall 18 to a large diameter at the cold inner partition wall 16. The large diameter at the one end may block the cooling flow more than the small diameter at the other end of the pedestal within the flow channel. This tapered arrangement may push the cooling flow towards the hot outer wall 18. The potential core flows may focus near the hot wall for cooling purposes. [0047] In operation, the temperature of the inner partition wall 16 can be colder than the temperature of the outer wall 18 due to the drop off of heat transfer performance. When the cooling air 20 flows around the plurality of pedestals 26 the cooling air 20 may be tripped or disturbed by the discrete trip strips 30 built around the plurality of pedestals 26. A high level of turbulent flow may thus be generated. The tripping of the cooling air 20 may eliminate the drop off of heat transfer performance for the pedestal. The highly turbulent flow, versus without the pedestal, and the increased internal convective area provided by the pedestal and discrete trip strips 30 may result in better heat transfer performance than that obtained from current pedestal cooling designs. In addition, the pedestals 26 may block-off some of the channel flow area and increase the channel through flow velocity. While the second set of discrete trip strips 32 on the inner wall of the airfoil may trip the boundary layer. A much higher heat transfer coefficient for the cooling channel may be generated with a corresponding decrease in temperature that may increase the life of the component.

[0048] 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.

Claims

CLAIMS What is claimed is:

1. An air cooled turbine airfoil comprising:

a leading edge (12) and a trailing edge (14);

an outer wall (18) comprising a pressure side wall and a suction side wall extending between the leading edge (12) and the trailing edge (14);

a multiple pass serpentine flow cooling circuit extending along the turbine airfoil from the leading edge (12) region to the trailing edge (14) region; and

a plurality of pedestals (26) extending across at least one of the channels of the serpentine flow cooling circuit, the plurality of pedestals (26) spread apart in a chordal direction (C) to define radial passages for aft flowing cooling air (20),

a plurality of discrete trip strips (30) incorporated along the sides of each of the plurality of pedestals,

wherein at least a portion of the quantity of plurality of pedestals (26) are tapered in length from a large diameter end (38) to a small diameter end (40),

wherein the small diameter end (40) of the plurality of pedestals (26) is adjacent to the outer wall (18) of the turbine airfoil.

2. The air cooled turbine airfoil according to claim 1, wherein the plurality of pedestals (26) are each in a conical shape.

3. The air cooled turbine airfoil according to any of the claims 1 through 2, further comprising a chord-wise extending inner partition wall (16) substantially running the length of the airfoil, wherein the plurality of pedestals (26) are disposed so that the large diameter end (38) is adjacent to the inner partition wall (16) and the smaller diameter end is adjacent to the outer wall (18) of the airfoil.

4. The air cooled turbine airfoil according to any of the claims 1 through 3, further comprising a second set of discrete trip strips (32) that are connected to an airfoil inner channel wall (36) in between the plurality of pedestals (26).

5. The air cooled turbine airfoil according to any of the claims 1 through 4, wherein the airfoil and plurality of pedestals (26) are formed as a single piece.

6. The air cooled turbine airfoil according to any of the claims 1 through 4, wherein the airfoil and plurality of pedestals (26) are assembled through multiple ceramic core pieces to form an integral ceramic core.

7. The air cooled turbine airfoil according to any of the claims 1 through

6, wherein the airfoil and plurality of pedestals (26) are printed using a high temperature resistant material.

8. The air cooled turbine airfoil according to any of the claims 3 through

7, wherein six discrete trip strips (30) are arranged at approximately 60 degrees from each other along each of the plurality of pedestals (26).

9. The air cooled turbine airfoil according to any of the claims 1 through 7, wherein four discrete trip strips (30) are arranged at approximately 90 degrees from each other along each of the plurality of pedestals (26).

10. A method for increasing a heat transfer coefficient through a cooling circuit of a turbine blade (10), comprising:

providing a turbine airfoil comprising:

a leading edge (12) and a trailing edge (14);

an outer wall (18) comprising a pressure side wall and a suction side wall extending between the leading edge (12) and trailing edge

(14);

a plurality of pedestals (26) extending across at least one of the channels of the serpentine flow cooling circuit,

a plurality of discrete trip strips (30) positioned along the sides of each pedestal, wherein at least a portion of the plurality of pedestals (26) are tapered in length from a large diameter end (38) to a small diameter end (40),

wherein the small diameter end (40) of the plurality of pedestals (26) is adjacent to the outer wall (18) of the turbine airfoil; and

sending cooling air (20) through the multiple pass serpentine flow cooling circuit,

passing the cooling air 20 around the plurality of pedestals (26), wherein the tapered shape of the plurality of pedestals (26) forces the cooling air 20 towards the small diameter end (40) with an increased heat transfer coefficient.

1 1. The method according to claim 10, wherein the plurality of pedestals (26) are each in a conical shape.

12. The method according to any of claims 10 or 1 1, wherein the airfoil further comprises a chord-wise extending inner partition wall (16) substantially running the length of the airfoil, wherein the plurality of pedestals (26) are disposed so that the large diameter end (38) is adjacent to the inner partition wall (16) and a smaller diameter is adjacent to the outer wall (18) of the airfoil.

13. The method according to any of claims 10 through 12, wherein the turbine airfoil further comprises a second set of discrete trip strips (32) that are connected to an airfoil inner channel wall (36) in between the plurality of pedestals (26).