Patent Application
20100186415
This invention relates to internal cooling within a gas turbine engine; and more particularly, to an assembly and method for providing better and more uniform cooling in a transition region between a combustor liner and a transition duct that directs combustion gases to the first stage of the turbine.
Traditional gas turbine combustors use diffusion (i.e., non-premixed) combustion in which fuel and air enter the combustion chamber separately. The process of mixing and burning produces flame temperatures exceeding 3900° F. Since conventional combustor liners and/or transition pieces are generally capable of withstanding a maximum temperature of only about 1500° F. (about 820° C.) for about ten thousand hours (10,000 hrs), steps to protect the combustor liner and/or transition duct, as well as the seal construction at the interface of the combustor liner and transition piece, must be taken for durability, creep resistance and seal integrity. This has typically been done by film-cooling which involves introducing relatively cool compressor air into a plenum formed by the combustor liner surrounding the outside of the combustor. In this prior arrangement, the air from the plenum passes through louvers in the combustor liner and then passes as a film over the inner surface of the liner, thereby maintaining combustor liner integrity.
Because diatomic nitrogen rapidly disassociates at temperatures exceeding about 3000° F. (about 1650° C.), the high temperatures of diffusion combustion result in relatively large NOx emissions. One approach to reducing NOx emissions has been to premix the maximum possible amount of compressor air with fuel. The resulting lean premixed combustion produces cooler flame temperatures and thus lower NOx emissions. NOx emissions reduction through premixed combustion is limited by the fraction of total compressor air available for combustion. Although lean premixed combustion is cooler than diffusion combustion, the flame temperature is still too hot for prior conventional combustor components to withstand.
Furthermore, because the advanced combustors premix the maximum possible amount of air with the fuel for NOx reduction, little or no cooling air is available, making film-cooling of the combustor liner and transition piece impractical. Nevertheless, combustor liners require active cooling to maintain material temperatures below limits. In dry low NOx (DLN) emission systems, this cooling can only be supplied as cold side convection. Such cooling must be performed within the requirements of thermal gradients and pressure loss. Thus, means such as thermal barrier coatings in conjunction with “backside” cooling have been considered to protect the combustor liner and transition piece from damage by such high heat. Backside cooling involved passing the compressor discharge air over the outer surface of the transition piece and combustor liner prior to premixing the air with the fuel.
Another current practice is to impingement cool the liner, and, optionally, to provide turbulators on the exterior surface of the liner (see, for example, U.S. Pat. No. 7,010,921). Still another practice is to provide an array of concavities on the exterior or outside surface of the liner (see U.S. Pat. No. 6,098,397). These various known techniques enhance heat transfer but with varying effects on thermal gradients and pressure losses.
Another technique, as described in commonly owned U.S. Pat. No. 7,010,921, provides straight axial cooling air channels, radially between the liner and the seal at the aft end of the liner, designed especially to cool the seal.
There remains a need, however to provide even more effective cooling in the combustor liner/transition piece interface region to further increase the durability and hence useful life of the combustor liners and associated seals.
The above discussed and other drawbacks and deficiencies are at least partially overcome or alleviated in an example embodiment by an apparatus for cooling the interface region between the combustor liner and the transition piece of a gas turbine.
Thus, in one aspect, the invention relates to a combustor liner comprising an open-ended, generally cylindrical body having a forward end and an aft end, the aft end formed with a plurality of axially extending channels defined by a plurality of axially extending, circumferentially spaced ribs; each channel provided with a plurality of axially-spaced transverse turbulators, the ribs having a height greater than the turbulators.
In another aspect, the invention relates to a combustor for a turbine comprising: a combustor liner; a first flow sleeve surrounding the combustor liner with a first flow annulus therebetween, the first flow sleeve having a plurality of cooling apertures formed about a circumference thereof for directing compressor discharge air into the first flow annulus; a transition piece body connected to the combustor liner, the transition piece body being adapted to carry hot combustion gases to the turbine; a second flow sleeve surrounding the transition piece body, the second flow sleeve having a second plurality of cooling apertures for directing compressor discharge air into a second flow annulus between the second flow sleeve and the transition piece body, the first flow annulus connecting to the second flow annulus; a resilient seal structure disposed radially between an aft end portion of the combustor liner and a forward end portion of the transition piece body; a cover sleeve disposed radially between the aft end portion of the combustor liner and the resilient seal structure, a plurality of axially-extending, circumferentially-spaced air flow channels between the cover sleeve and the aft end portion of the combustor liner; and a plurality of axially-spaced, transversely-oriented turbulators in each of the air flow channels, projecting towards but spaced from the cover sleeve.
In still another embodiment, the invention relates to a method of cooling a transition region in a gas turbine combustor between an aft end portion of a combustor liner and a forward end portion of a transition piece, the combustor liner having a first flow sleeve surrounding the combustor liner with a first flow annulus therebetween, the first flow sleeve having a first plurality of cooling apertures formed about a circumference thereof for directing compressor discharge air into the first flow annulus, the transition piece connected to the combustor liner and adapted to carry hot combustion gases to the turbine; a second flow sleeve surrounding the transition piece, the second flow sleeve having a second plurality of cooling apertures for directing compressor discharge air into a second flow annulus between the second flow sleeve and the transition piece, the first flow annulus connecting to the second flow annulus; the transition region including a resilient seal structure disposed radially between the aft end portion of the combustor liner and the forward end portion of the transition piece; the method comprising: (a) configuring the aft end portion of the combustor liner to include a plurality of axially oriented flow channels, and a plurality of radially outwardly projecting, transverse turbulators in each of the flow channels; (b) disposing a cover sleeve between the aft end portion of the combustor liner and the resilient seal structure so as to close a radially outer side of the flow channels; the transverse turbulators projecting towards but being spaced from the cover sleeve; and (c) supplying compressor discharge air through at least some of the first and second pluralities of cooling apertures and through the flow channels to thereby cool the resilient seal.
The invention will now be described in greater detail in conjunction with the drawings identified below.
Flow from the gas turbine compressor (not shown) enters into a case 24. About 50% of the compressor discharge air passes through apertures (not shown in detail) formed along and about the transition piece impingement sleeve 16 for flow in an annular region or annulus 26 between the transition piece body 14 and the radially outer transition piece impingement sleeve 16. The remaining approximately 50% of the compressor discharge flow passes into flow sleeve holes 28 of the upstream combustion liner flow sleeve 20 and into an annulus 30 between the flow sleeve 20 and the liner 18 and eventually mixes with the air from the downstream annulus 26. The combined air eventually mixes with the gas turbine fuel in the combustion chamber.
As previously noted, the hot gas temperature at the aft end of the liner 18, and the connection or interface region 22, is approximately 2800° F. However, the liner metal temperature at the downstream, outlet portion of interface region 22 is preferably about 1400-1550° F. As described in greater detail below, to help cool the liner 18 to this lower metal temperature range during passage of heated gases through the interface region 22, the aft end of the liner 18 has been formed with axial passages through which cooling air is flowed. This cooling air serves to draw off heat from the liner and thereby significantly lower the liner metal temperature relative to that of the hot gases.
More specifically, and as best seen in
In accordance with an exemplary but nonlimiting embodiment of this invention, the cooling arrangement shown in
One analysis conducted to date shows temperature reductions of 50-100° F. in the interface region. Therefore, by providing the transverse turbulators 52 as proposed herein, it should be possible to achieve greater heat transfer with the same amount of cooling air (or the same amount of heat transfer with less cooling air), as compared to non-turbulated flow channels. This additional cooling capability increases service life and/or the ability to fire the gas turbine at higher temperatures and/or enables reduced NOx emissions.
While the invention has been described in connection with what is presently considered to be the most practical and preferred embodiment, it is to be understood that the invention is not to be limited to the disclosed embodiment, but on the contrary, is intended to cover various modifications and equivalent arrangements included within the spirit and scope of the appended claims.
