This invention relates to turbomachinery and specifically, to the cooling of combustor and transition pieces in gas turbine combustors.
Conventional gas turbine combustion systems employ multiple combustor assemblies to achieve reliable and efficient turbine operation. Each combustor assembly includes a cylindrical liner, a fuel injection system, and a transition piece that guides the flow of the hot gases from the combustor to the inlet of the turbine. Generally, a portion of the compressor discharge air is used to cool the combustor liner and is then introduced into the combustor reaction zone to be mixed with the fuel and burned.
In systems incorporating impingement cooled transition pieces, a hollow flow sleeve surrounds the transition piece, and the flow sleeve wall is perforated so that compressor discharge air will flow through the cooling apertures in the sleeve wall and impinge upon (and thus cool) the transition piece. This cooling air then flows along an annulus between the flow sleeve and the transition piece, and then into another annulus between the combustor liner and a second flow sleeve surrounding the liner. The second flow sleeve is also formed with several rows of cooling holes about its circumference, the first row located adjacent a mounting flange where the second flow sleeve joins to the first flow sleeve.
In combustor configurations utilizing impingement cooling for the combustor liner and/or transition piece (or other combustor component), it is often the case that the pitch between adjacent impingement jets tends to be too large to effectively cool the component. Specifically, the large pitch spacing gives rise to areas which are left uncooled (sometimes referred to as “hot spots), and also to excessive thermal gradients. There remains a need therefore to improve the cooling efficiency of impingement cooled combustor components.
In accordance with exemplary but nonlimiting embodiments, this invention employs effusion cooling in regions where impingement cooling is deficient. Thus, in one aspect, the present invention relates to a cooling arrangement for a first turbine combustor component surrounded by a second turbine combustor component, the cooling arrangement comprising: a first plurality of impingement cooling holes in the second turbine combustor component, the plurality of impingement cooling holes directing cooling air onto designated areas of the first turbine combustor component; and a second plurality of effusion cooling holes in the first turbine combustor component located to cool by effusion other areas of the first turbine combustor component.
In another aspect, the invention relates to a method of cooling a turbine combustor component comprising: (a) surrounding the turbine combustor component with a flow sleeve, with an annular flow passage between the turbine component and the flow sleeve; (b). providing a plurality of impingement cooling holes in the flow sleeve adapted to supply cooling air onto designated areas of the turbine component; and (c) providing a plurality of effusion cooling holes in the turbine combustor component adapted to supply cooling air to other designated areas of the turbine combustor component.
The invention will now be described in detail in connection with the drawings identified below.
Referring to
More specifically, in an exemplary but nonlimiting embodiment, the compressor discharge air flows through an annular gap 30 formed by a first flow sleeve 32 surrounding the transition piece 20 and a second flow sleeve 34 surrounding the liner 24. Each flow sleeve 32, 34 has a series of holes, slots, or other openings (not shown, but see similar holes in
In the exemplary but nonlimiting embodiment shown in
Because of the typical large pitch spacing between adjacent impingement hole cooling jets, however, liner cooling is less than optimal. To supplement and enhance the impingement cooling, effusion cooling apertures 44 have been added to the liner 46. More specifically, one or more arrays 48 of effusion cooling apertures 44 are formed in the liner 46 in selected locations where impingement cooling in insufficient.
As shown in
In an exemplary but nonlimiting implementation, the impingement holes may have diameters in the range of from about b 0.10 to about 1.0 in. (or if noncircular, substantially equivalent cross-sectional areas). The smaller effusion holes may have diameters in the range of from about 0.02 to about 0.04 in. (or if noncircular, substantially equivalent cross-sectional areas).
The combination of impingement and effusion cooling may be applied to any component where impingement jet pitch spacing yields unfavorable thermal conditions. Such components include but are not limited to combustor liners and transition ducts (or pieces) that supply the hot combustion gases to the first stage nozzle. The number, size, shape and pattern(s) of the impingement cooling holes and the effusion cooling holes are not intended to be limited in any way.
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.