The present invention is directed to hot-section turbine components in gas turbine engines, and more specifically to cooling structures of turbine components.
A gas turbine engine includes, in sequential flow order: a compressor, a combustor, and one or more turbines. During operation, air is compressed through the compressor and is mixed with fuel in the combustor. The fuel is ignited in the combustor, generating hot, high energy combustion gases which flow downstream through the turbine stages. The turbine stages extract energy from these combustion gases.
To prolong the service life of turbine blades and other hot section components and to reduce engine operating cost, portions of these components, for example turbine blade tips, often employ “active cooling”. This type of cooling is effected by bleeding off pressurized air at a relatively low temperature from some other portion of the engine such as the compressor, and then discharging the bleed air through cooling holes to form a protective film.
One problem with active cooling is that the use of bleed air is expensive in terms of overall fuel consumption.
Another problem with active cooling, particularly for turbine blade tips, is that the cooling holes can be damaged during a “rub” event in which the blade tips contact the surrounding shroud, thus lowering cooling effectiveness.
At least one of these problems is addressed by a turbine component cooling structure in which transpiration cooling is effected using air cavities that form an organized matrix of torturous passages. This structure may incorporate a pattern of recessed areas in a blade tip that makes the cooling pattern robust so that it does not get closed by material sacrificed during a rub.
According to one aspect of the technology described herein, a tip cooling apparatus for a turbine airfoil includes: a tip cap; a pair of spaced-apart tip walls connected to, extending around, and projecting outwardly from the tip cap so as to surround a central portion of the tip cap; a pocket defined by the tip walls; at least one feed hole passing through the tip cap or tip walls, communicating with the pocket; and a cooling matrix disposed in the pocket, the cooling matrix being an organized structure including an inlet surface having a plurality of inlets communicating with the pocket, and an outlet surface having a plurality of outlets, and further comprising a plurality of interior passages interconnecting the inlets to the outlets, with no line-of-sight therebetween.
According to another aspect of the technology described herein, a turbine blade includes: a dovetail; a blade shank extending from the dovetail and terminating in a platform that projects laterally outwardly from and surrounds the shank; and an airfoil extending from the platform, the airfoil including: a root, a tip, a leading edge, a trailing edge, a pressure side wall and a suction side wall, and an internal cavity, said internal cavity being bounded at its radially outer end by a tip cap; a squealer tip comprising spaced-apart tip walls extending above said tip cap; a cooling matrix disposed outboard of the tip cap, the cooling matrix being an organized structure including an inlet surface having a plurality of inlets communicating with the interior cavity, and an outlet surface having a plurality of outlets, and further comprising a plurality of interior passages interconnecting the inlets to the outlets, with no line-of-sight therebetween.
According to another aspect of the technology described herein, a turbine component includes: a flowpath wall defining a flowpath surface configured to be exposed in operation to a hot gas flow, the flowpath wall having a backside opposite to the flowpath surface which forms part of the boundary of an internal cooling circuit of the component; a pocket defined within the flowpath wall; at least one feed hole formed in the flowpath wall communicating with the pocket; and a cooling matrix disposed in the pocket, the cooling matrix being an organized structure including an inlet surface having a plurality of inlets communicating with the pocket, and an outlet surface having a plurality of outlets, and further comprising a plurality of interior passages interconnecting the inlets to the outlets, with no line-of-sight therebetween.
The invention may be best understood by reference to the following description taken in conjunction with the accompanying drawing figures in which:
Referring to the drawings wherein identical reference numerals denote the same elements throughout the various views,
The airfoil 18 includes an internal cooling circuit (not shown) which may have any conventional configuration, such as a serpentine circuit. The cooling circuit extends through the platform 16 and dovetail 12 and includes inlets in the base of the dovetail 12 for receiving pressurized cooling air from a compressor of the gas turbine engine (not shown) in a conventional manner. In this way, the airfoil 18 is internally cooled by the cooling air which then may be discharged through the thin airfoil sidewalls in various rows of film cooling holes 35 of conventional size and configuration. The airfoil 18 may incorporate a plurality of trailing edge cooling holes 32, or it may incorporate a number of trailing edge bleed slots (not shown) on the pressure side wall 24 of the airfoil 18.
An upstanding squealer tip 36 extends radially outwardly from the tip cap 34 and is disposed in close proximity to a stationary shroud (not shown) in the assembled engine, in order to minimize airflow losses past the tip 22. The tip structure 36 comprises a suction side tip wall 38 disposed in a spaced-apart relationship to a pressure side tip wall 40. The tip walls 40 and 38 form extensions of the pressure and suction side walls 24 and 26, respectively, and may be integral to the airfoil 18. The outer surfaces of the pressure and suction side tip walls 40 and 38 respectively form continuous surfaces with the outer surfaces of the pressure and suction side walls 24 and 26. Radial end faces of the tip walls 40 and 38 define a tip surface 44. According to the principles described herein, the squealer tip 36 may incorporate a cooling matrix configured to provide transpiration cooling of the airfoil tip 22 in operation.
In the example shown in
The cooling matrix 48 is an organized (i.e. not random) structure defining a tortuous air path extending from a plurality of inlets 58 at the inlet surface (see
At least some of the tubes 62 are configured with their centerlines 66 extending at a non-parallel angle to a radial direction “R” of the airfoil 18. The tubes 62 may also be configured such that at least some of the tubes intersecting a given one of the hubs 64 are not aligned coaxially with each other. Stated another way, the central axes of two tubes meeting at a common hub 64 may be arranged so they form an acute angle. The combination of these features guarantees that no line-of-sight exists along the radial direction R.
As shown in
In another example shown in
The operation of the cooling matrix 48 may be understood with reference to
In operation, it is possible that the rotating blade tip 22 may contact the surrounding shroud, an event known as a “rub”. During a rub, the forces involved may tend to have an effect of laterally displacing or “smearing” material at the tip of the blade. In a conventional turbine airfoil using transpiration film cooling holes in the tip, the laterally-displaced material can close off the transpiration film cooling resulting in a loss of cooling effectiveness.
To counteract this effect and provide continued cooling effectiveness despite rubs, the cooling matrix 48 may configured to be rub-resistant, as illustrated in
All or part of the tip structure described above, including the tip walls and the cooling matrix, or portions thereof, may be part of a single unitary, one-piece, or monolithic component, and may be manufactured using a manufacturing process which involves layer-by-layer construction or additive fabrication (as opposed to material removal as with conventional machining processes). Such processes may be referred to as “rapid manufacturing processes” and/or “additive manufacturing processes,” with the term “additive manufacturing process” being the term used herein to refer generally to such processes. Additive manufacturing processes include, but are not limited to: Direct Metal Laser Melting (DMLM), Laser Net Shape Manufacturing (LNSM), electron beam sintering, Selective Laser Sintering (SLS), 3D printing, such as by inkjets and laserjets, Stereolithography (SLS), Electron Beam Melting (EBM), Laser Engineered Net Shaping (LENS), and Direct Metal Deposition (DMD). It is believed that the cooling matrix configurations described above cannot be created, or could not be practically manufactured, using a conventional (non-additive) method.
In one example, the cooling matrix may be formed using an additive manufacturing process, and then installed into the pockets described above, for example using a conventional brazing process.
In another example, additive manufacturing processes may be used to form the cooling matrix described herein as well as solid structure in the same monolithic element. For example, certain elements have been described as having solid walls forming a pocket with a cooling matrix disposed therein. It will be understood that an additive manufacturing process could be used to build both of those elements simultaneously, with the portion designated as being the cooling matrix incorporating the internal passages described above, and the portions described as solid structure lacking internal passages.
The tip structure described herein has advantages over the prior art. The cooling configuration is more effective than conventional convection or conduction cooling. First, it creates transpiration film on the tip cap surface by creating a high pressure reduction. Second, it creates a torturous path in the organized matrix of air passages that induces a large heat transfer coefficient.
Furthermore, the optional pattern of recessed areas makes the tip robust to rubs by recessing part of the air orifices below the rub height allowed for on the tip.
Cooling the tip effectively will increase field life of turbine blades, and reduce costs to the customer. Additionally, tip life is important to blade clearances which in turn affect engine performance.
The organized cooling matrix 48 described herein may be used to cool elements other than airfoil tips. A gas turbine engine includes numerous hot section components having a surface exposed to a hot gas flow, herein referred to as a “flowpath wall”, and any such flowpath wall could incorporate an organized cooling matrix. For example, the platform 16 of the turbine blade 10, as well as its pressure side wall 24 and suction side wall 26, constitute flowpath walls. As another example,
In operation, the cooling circuit 255 would be supplied with pressurized cooling air as described above for the turbine blade 10. The function of the cooling matrix 248 is substantially as described above for the cooling matrix 48.
The foregoing has described a cooling structure for a gas turbine engine component. All of the features disclosed in this specification (including any accompanying claims, abstract and drawings), and/or all of the steps of any method or process so disclosed, may be combined in any combination, except combinations where at least some of such features and/or steps are mutually exclusive.
Each feature disclosed in this specification (including any accompanying claims, abstract and drawings) may be replaced by alternative features serving the same, equivalent or similar purpose, unless expressly stated otherwise. Thus, unless expressly stated otherwise, each feature disclosed is one example only of a generic series of equivalent or similar features.
The invention is not restricted to the details of the foregoing embodiment(s). The invention extends to any novel one, or any novel combination, of the features disclosed in this specification (including any accompanying claims, abstract and drawings), or to any novel one, or any novel combination, of the steps of any method or process so disclosed.
Number | Name | Date | Kind |
---|---|---|---|
4487550 | Horvath | Dec 1984 | A |
5660523 | Lee | Aug 1997 | A |
5738491 | Lee et al. | Apr 1998 | A |
6135715 | Correia | Oct 2000 | A |
6155778 | Lee | Dec 2000 | A |
6461107 | Lee | Oct 2002 | B1 |
7402335 | Bolms | Jul 2008 | B2 |
7413001 | Wang et al. | Aug 2008 | B2 |
7537431 | Liang | May 2009 | B1 |
7997865 | Liang | Aug 2011 | B1 |
8061987 | Liang | Nov 2011 | B1 |
8801377 | Liang | Aug 2014 | B1 |
9039917 | Szuromi et al. | May 2015 | B2 |
9249670 | Bunker | Feb 2016 | B2 |
9297262 | Zhang et al. | Mar 2016 | B2 |
10036258 | Mongillo | Jul 2018 | B2 |
20100119377 | Tibbott | May 2010 | A1 |
20110262695 | Lee et al. | Oct 2011 | A1 |
20150068629 | Kottilingam et al. | Mar 2015 | A1 |
20150144496 | Morris et al. | May 2015 | A1 |
20150330228 | Quach | Nov 2015 | A1 |
20150345304 | Mongillo et al. | Dec 2015 | A1 |
20160090848 | Engeli et al. | Mar 2016 | A1 |
20160312624 | Loricco et al. | Oct 2016 | A1 |
Number | Date | Country |
---|---|---|
2004202946 | Feb 2005 | AU |
2469030 | Jun 2012 | EP |
2279705 | Jan 1995 | GB |
Entry |
---|
International Search Report and Written Opinion issued in connection with corresponding PCT Application No. PCT/US2017/060021 dated Jul. 26, 2018. |
Number | Date | Country | |
---|---|---|---|
20180142559 A1 | May 2018 | US |