Abstract:
A combustor assembly for a turbine includes a combustor and a combustor liner; a first flow sleeve surrounding the combustor liner forming a first substantially axially-extending flow annulus radially therebetween. The first flow sleeve has a first plurality of apertures formed about a circumference thereof for directing compressor discharge air as cooling air radially into said first flow annulus. A transition piece is connected to the combustor liner, the transition piece adapted to carry hot combustion gases to the turbine, and a second flow sleeve surrounds the transition piece forming a second substantially axially-extending flow annulus radially therebetween. The second flow sleeve has a second plurality of apertures for directing compressor discharge air as cooling air radially into the second flow annulus, the first substantially axially-extending flow annulus connecting with the second substantially axially-extending flow annulus. A resilient annular seal structure is disposed radially between an aft end portion of the combustor liner and a forward end portion of said transition piece, the resilient annular seal structure configured to form a first annular cavity radially between the forward end portion of the transition piece and the aft end portion of said combustor liner. At least one transfer tube radially extends from the second flow sleeve through the second flow annulus to the transition piece, and is arranged to supply compressor discharge cooling air radially from an area outside the first and second substantially axially extending flow annuli directly to the resilient annular seal structure and to the aft end of the combustor liner.

Description:
BACKGROUND OF THE INVENTION 
       [0001]    This invention relates to internal cooling within a gas turbine engine, and more particularly, to an assembly for providing more efficient and uniform cooling in an interface or transition region between a combustor liner and a transition duct. 
         [0002]    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 combustors and/or transition pieces (or ducts) having liners are generally capable of withstanding a maximum temperature on the order of only about 1500° F. for about ten thousand hours (10,000 hrs), steps to protect the combustor and/or transition piece must be taken. Typically, this has been done by a combination of impingement and film-cooling which involves introducing relatively cool compressor discharge air into a plenum formed by a flow sleeve surrounding the outside of the combustor liner. In this prior arrangement, the air from the plenum passes through apertures in the combustor liner and impinges on the exterior liner surface and then passes as a film over the outer or cold-side surface of the liner. 
         [0003]    Because advanced combustors premix the maximum possible amount of air with the fuel for NOx reduction, however, little or no cooling air is available, thereby making film-cooling of the combustor liner and transition piece problematic. 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 due to excessive heat. Backside cooling involves passing the compressor discharge air over the outer surface of the transition piece and combustor liner prior to premixing the air with the fuel. 
         [0004]    With respect to the combustor liner, another current practice is to impingement cool the liner, or to provide turbulators on the exterior surface of the liner (see, for example, U.S. Pat. No. 7,010,921). Turbulation works by providing a blunt body in the flow which disrupts the flow creating shear layers and high turbulence to enhance heat transfer on the surface. Another practice is to provide an array of concavities on the exterior or outside surface of the liner (see, for example, U.S. Pat. No. 6,098,397). Dimple concavities function by providing organized vortices that enhance flow mixing and scrub the surface to improve heat transfer. The various known techniques enhance heat transfer but with varying effects on thermal gradients and pressure losses. 
         [0005]    There remains a need for more efficient and more uniform cooling at the combustor liner/transition piece seal interface, and for minimizing leakage at the interface seal where cooling air is routed to the seal region from a higher-pressure location for the purpose of cooling the seal and adjourning components. 
       BRIEF DESCRIPTION OF THE INVENTION 
       [0006]    The above-mentioned drawbacks (and others) are overcome or alleviated in example embodiments as broadly described below. 
         [0007]    Thus, in one exemplary but nonlimiting embodiment, there is provided a combustor assembly for a turbine comprising a combustor including a combustor liner; a first flow sleeve surrounding the combustor liner forming a first substantially axially-extending flow annulus radially therebetween, the first flow sleeve having a first plurality of apertures formed about a circumference thereof for directing compressor discharge air as cooling air radially into the first flow annulus; a transition piece connected to the combustor liner, the transition piece adapted to carry hot combustion gases to the turbine; a second flow sleeve surrounding the transition piece forming a second substantially axially-extending flow annulus radially therebetween, the second flow sleeve having a second plurality of apertures for directing compressor discharge air as cooling air radially into the second flow annulus, the first substantially axially-extending flow annulus connecting with the second substantially axially-extending flow annulus; a resilient annular seal structure disposed radially between an aft end portion of the combustor liner and a forward end portion of the transition piece, the resilient annular seal structure configured to form a first annular cavity radially between the forward end portion of the transition piece and the aft end portion of the combustor liner; and at least one transfer tube radially extending from the second flow sleeve through the second flow annulus to the transition piece, and arranged to supply compressor discharge cooling air radially from an area outside the first and second substantially axially-extending flow annuli directly to the resilient annular seal structure and to the aft end of the combustor liner. 
         [0008]    In another exemplary but nonlimiting aspect, there is provided a combustor assembly for a turbine comprising a combustor including a combustor liner; a first flow sleeve surrounding the combustor liner forming a first substantially axially-extending flow annulus radially therebetween, the first flow sleeve having a first plurality of apertures formed about a circumference thereof for directing compressor discharge air as cooling air radially into the first flow annulus; a transition piece connected to the combustor liner, the transition piece adapted to carry hot combustion gases to the turbine; a second flow sleeve surrounding the transition piece forming a second substantially axially-extending flow annulus radially therebetween, the second flow sleeve having a second plurality of apertures for directing compressor discharge air as cooling air radially into the second flow annulus, the first substantially axially-extending flow annulus connecting with the second substantially axially-extending flow annulus; a resilient annular seal structure disposed radially between an aft end portion of the combustor liner and a forward end portion of the transition piece; and means for supplying compressor discharge cooling air from a location external to the first and second flow sleeves directly to the resilient annular seal structure and an aft end portion of the combustor liner. 
         [0009]    In still another exemplary but nonlimiting embodiment, there is provided a method of cooling an aft end portion of a gas turbine combustor liner and an annular seal structure radially interposed between the aft end portion of the gas turbine combustor liner and a transition piece adapted to supply combustion gases from the combustor liner to a first stage of the gas turbine, and wherein the combustor liner is connected to the transition piece, and a flow sleeve surrounding the combustor liner is connected to an impingement sleeve surrounding the transition piece thereby forming a cooling flow annulus, the method comprising supplying cooling air from a location external to the flow sleeve and the impingement sleeve directly to the annular seal structure and the aft end portion of the combustor liner; and thereafter directing at least a major portion of the cooling air into the cooling flow annulus. 
         [0010]    The invention will now be disclosed in detail in connection with the drawings identified below. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0011]      FIG. 1  is a partial schematic illustration of a gas turbine combustor section including a combustor liner/transition piece interface region; 
           [0012]      FIG. 2  is a partial but more detailed perspective of a combustor liner and flow sleeve joined to a transition piece and impingement sleeve with an annular seal located between the transition piece and combustor liner; 
           [0013]      FIG. 3  is an exploded partial view, of the aft end of a conventional combustion liner illustrating a cooling arrangement for a combustor liner-transition piece hula seal; 
           [0014]      FIG. 4  is a partial perspective view, partially cut away, illustrating a cooling arrangement for a hula seal in accordance with an exemplary but nonlimiting embodiment of the invention; 
           [0015]      FIG. 5  is a cross-sectional elevational view of the arrangement shown in  FIG. 4 ; 
           [0016]      FIG. 6  is a simplified, partial section of a cooling arrangement in accordance with a second exemplary but nonlimiting embodiment; 
           [0017]      FIG. 7  is a simplified, partial section of a third cooling arrangement in accordance with another exemplary but nonlimiting embodiment; 
           [0018]      FIG. 7A  is a cross section taken along the line  7 A- 7 A in  FIG. 7 ; 
           [0019]      FIG. 8  is a simplified, partial section of a fourth cooling arrangement in accordance with another exemplary but nonlimiting embodiment; 
           [0020]      FIG. 8A  is a partial section taken along the line  8 A- 8 A in  FIG. 8 ; 
           [0021]      FIG. 9  is a simplified, partial section of a fifth cooling arrangement in accordance with another exemplary but nonlimiting embodiment; 
           [0022]      FIG. 10  is a simplified, partial section of a sixth cooling arrangement in accordance with another exemplary but nonlimiting embodiment; 
           [0023]      FIG. 11  is a simplified, partial section of a seventh cooling arrangement in accordance with another exemplary but nonlimiting embodiment; and 
           [0024]      FIG. 12  is a simplified, partial section of an eighth cooling arrangement in accordance with another exemplary but nonlimiting embodiment. 
       
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       [0025]      FIG. 1  schematically depicts the aft end of a turbine combustor  10  and its connection to a transition piece or duct assembly  12  that directs the hot combustion gases to the first stage of the turbine. The transition piece assembly  12  includes a radially inner transition piece body (or simply, transition piece)  14  and an impingement sleeve (or second flow sleeve)  16  spaced radially outward of the transition piece  14 . Upstream thereof (relative to the flow of combustion gases from the combustor to the turbine first stage, indicated by flow arrows CG) is the radially inner combustion liner  18  and its associated radially outer flow sleeve (or first flow sleeve)  20 . The encircled region  22  is the transition piece/combustor liner interface that is of interest. 
         [0026]    Flow from the gas turbine compressor (not shown) enters into the turbine or machine casing  24  as indicated by flow arrows F. About 50% of the so-called compressor discharge air passes radially through apertures (not shown in detail) formed along and about the impingement sleeve  16  as indicated by flow arrows CD. This air is reverse-flowed (i.e., toward the forward end of the combustor, counter to the flow of gases within the combustor liner and transition piece) in an annular region or passage  26  between the transition piece  14  and the impingement sleeve  16 . The remaining approximately 50% of the compressor discharge air passes into holes  28  in the flow sleeve  20  and into an annular passage  30  between the flow sleeve  20  and the liner  18 , where it mixes with the air flowing in the annular passage  26 . The combined air from passages  26  and  30 , used initially to cool the transition piece and combustor liner, eventually reverses direction again before entering the combustor liner where it mixes with the gas turbine fuel for burning in the combustion chamber  21 . 
         [0027]      FIG. 2  illustrates an exemplary connection at an interface  22  between the transition piece  14 /impingement sleeve  16 , and the combustor liner  18 /flow sleeve  20 . The impingement sleeve  16  is joined to a mounting flange  32  on the aft end of the flow sleeve  20 . Specifically, a radial outward piston seal  34  on the impingement sleeve  16  is received within a radially inward-facing annular groove  36  formed within the mounting flange  32 . The transition piece receives the combustor liner  18  in a telescoping relationship with a conventional, annular compression-type or hula seal  38  interposed therebetween. 
         [0028]    Referring now to  FIG. 3 , a prior cooling arrangement in the area of the interface hula seal  38  was designed to cool the aft end  50  of the combustor liner  18 . Specifically, the hula seal  38  is mounted radially between an annular cover plate  40  surrounding the liner aft end  50  and the transition piece  14  (see  FIG. 2 ). More specifically, the cover plate  40  forms a mounting surface for the compression or hula seal  38 . The aft end  50  of the liner  18  has a plurality of axial channels  42  formed by a plurality of axially-oriented raised sections or ribs  44  on the liner, closed on their radially outer sides by the plate  40 . Cooling air from the passage  26  is introduced into the channels  42  through air inlet apertures or openings  46  in the cover plate  40  at the forward end of the channels. The air then flows into and through the channels  42  and exits at the aft end  50  of the liner  18  to join the combustion gases flowing into the transition piece. See commonly-owned U.S. Pat. No. 7,010,921 for additional details. 
         [0029]      FIGS. 4 and 5  illustrate another combustor liner-transition piece interface that is similar in certain respects to those shown in  FIGS. 2 and 3  but with modifications as explained below in accordance with a first exemplary but nonlimiting example of the invention. 
         [0030]    In this first exemplary but nonlimiting embodiment, a transition piece  52  is connected to a combustor liner  54  at the aft end  56  of the liner. An impingement sleeve assembly  58  surrounds the transition piece  52  in radially-spaced relation thereto, forming a first annular flow passage  60 . A flow sleeve  62  surrounds the combustor liner  54 , also in radially spaced relation, thus forming a second annular flow passage  64  which is in direct flow communication with the first annular flow passage  60 . The impingement sleeve assembly  58  is joined to the substantially axial flow sleeve  62  by means of a radially outwardly directed annular piston seal  66  which is received in a radially inwardly facing groove  68  in an annular flange  70  at the aft end of the flow sleeve. The piston seal  66  is composed of a split, annular ring (similar to a piston ring), biased radially inwardly to maintain a minimum gap between the radially inner seal edge  61  and the forward end of the impingement sleeve assembly  58  (or, in the illustrated embodiment, the discrete coupling component  59  of the assembly  58 ). 
         [0031]    The aft end  56  of the combustor liner  54  may be formed with an annular array of substantially axially-oriented ribs  72  extending between an aft edge  74  of the liner and an annular shoulder or edge  76 , thus forming an array of axially-oriented channels  78  between respective rib pairs. The channels  78  are closed on their radially outer sides by an annular cover plate  80  that may be integral with or joined to (by welding, for example) the liner  54 . 
         [0032]    An annular row of cooling air exit holes  82  is provided at the forward end of the cover plate  80 , adjacent the annular shoulder  76 , and multiple annular rows or arrays of cooling air inlet holes  84  are provided nearer the aft end of the cover plate  80 . It will be appreciated that the arrangement and number of exit apertures or holes  82 ,  84  may be varied as required by specific cooling applications. 
         [0033]    A flexible, annular compression or hula seal  86  is telescoped over the aft end of the cover plate  80 , the seal comprising plural axially-extending and circumferentially-spaced spring fingers  88 , with axial slots  90  therebetween. 
         [0034]    The forward end  92  of the transition piece  52  is formed to include an annular plenum chamber  94  between radially outer and inner wall portions  96 ,  98 , respectively, of the transition piece body. Compressor discharge air external to the combustor (i.e. higher-pressure compressor air not flowing in the passages  60 ,  64 ) is supplied directly to the annular plenum chamber  94  by means of a plurality of circumferentially-spaced transfer tubes  100  extending radially between apertures  101  formed in the impingement sleeve assembly  58  and radially-aligned apertures  103  formed in the transition piece  52 . Note in this regard that the transfer tubes can be located within the discrete coupling component  59  of the transition piece assembly  58 . Absent a discrete coupling component, the transfer tubes would extend from apertures formed in the impingement sleeve itself. The transfer tubes  100  may be varied in number and may have various cross-sectional shapes including round, oval, oblong, airfoil, etc. 
         [0035]    Cooling air in the plenum  94  flows through circumferentially-spaced apertures  102  provided in the radially-inner wall portion  98  of the transition piece  52  and into an annular space or cavity  104  under the hula seal  86 , via the axial slots  90  between the spring fingers  88  of the seal. Depending on the arrangement of transfer tubes and their position relative to the hula seal spring  FIGS. 88 , the slots  90  may not be available for supplying air to the cavity  104 . In that case, discrete apertures  105  may be formed in the spring fingers  88 . The cooling air is now free to flow through the cooling holes  84  in the aft end of the cover plate  80  and into the channels  78 . Note, however, that the channels  78  are interrupted by one or more circumferentially extending ribs  106  located, in the exemplary embodiment, axially between the two rows of cooling holes  84  closer to the aft end of the hula seal  86  and the edge  74 . As a result, the cooling air will flow in two opposite directions on either side of the one or more ribs  106 . More specifically, the majority of the cooling air will flow toward the forward end of the combustor, exiting the apertures  82  and joining the air flowing in the passages  60 ,  64 , while a minor portion of the cooling air will flow toward the aft end of the combustor, exiting the channels  78  at edge  74  and joining the flow of combustion gases within the liner and transition duct. The major flow of cooling air thus cools the hula seal  86  and impingement cools the cold side of the aft end of the liner while the minor portion of the cooling air purges the seal cavity  104 , thus maintaining a flow of “fresh” cooling air through the cavity  104  and channels  78 . Here again, the number of transfer tubes  100  and the number of apertures  102  (total number and number per transfer tube) may vary as required by cooling requirements as well as combustor design requirements. It may also be advantageous in some circumstances to provide turbulators on the surfaces defining the channels  78  to enhance cooling. 
         [0036]    It will also be appreciated that by using discrete apertures  105  in the hula seal spring fingers  88 , the flow of cooling air into the space or cavity  104  can be better controlled than if the elongated slots  90  used as conduits for the supply of cooling air to the cavity  104 . Further in this regard, the apertures  105  may be sized and shaped to achieve optimum alignment with the apertures  102  when the components reach their maximum temperatures. 
         [0037]    Thus, by having the major portion of the cooling flow eventually join the flow in passage  64  to the combustor nozzle and having only a minor portion of the cooling flow purge the seal and escape into the combustion gas stream, seal leakage is minimized and air available for premixing (and hence reduced emissions) is increased while maintaining cooling efficiency. 
         [0038]      FIG. 6  represents an alternative exemplary but nonlimiting embodiment, illustrated in simplified form. As in the previously described embodiment, a liner  110  and flow sleeve  112  are joined to a transition duct  114  and its impingement sleeve  116  at an interface  118 . Circumferentially-spaced transfer tubes  120  extend radially between a coupling component  122  that joins the impingement sleeve  116  to the flow sleeve  112 , and the transition piece forward end  124 . In this embodiment, the hula seal  126  is inverted as compared to the arrangement in  FIGS. 4 and 5 , such that an annular space or cavity  128  is established radially outward of the seal  126 . Higher-pressure cooling air entering the annular cavity  128  via the transfer tubes  120  flows out of the annular space  128  via apertures  129  in the spring fingers (or through the slots between the spring fingers), in a direction toward the forward end of the combustor, joining the cooling flow in the passage  127  (corresponding to passage  64  in  FIGS. 4 and 5 ). Little to no cooling air escapes past the seal into the main combustion flow. In this embodiment, the seal  126  is impingement cooled and the interior cavity  128  is purged, but only marginal cooling of the aft end of the liner  110  is provided by convection cooling. 
         [0039]      FIGS. 7 and 7A  illustrate an embodiment similar to that shown in  FIGS. 4 and 5 . In this alternative design, there are no ribs as shown at  72  in  FIG. 4 , and hence no discrete channels  78 . Rather, a relatively smooth and continuous annular space or chamber  130  is formed radially between the aft end of the liner  132  and the annular cover plate  144 . In addition, the liner  132  is formed with an upturned aft edge  146 , defining in part the exit slots  148  for the minor portion of the purge air flowing through apertures  150  and the discrete annular chamber  152  (aft of the annular rib  156 ), subsequently exiting the slots  148  into the combustion gas stream. The major portion of cooling air flows through apertures  158  into the annular chamber  130  to impingement cool a portion of the aft end of the liner  132 , while convection cooling the adjacent upstream portion and subsequently exiting apertures  160  to join the flow of air between the combustor flow sleeve  163  and the liner  132 .  FIG. 7A  also illustrates a rounded, elongated cross-sectional shape for the transfer tube  162 . Aside from these differences, the arrangement is otherwise substantially as shown and described above in connection with  FIGS. 4 and 5 . The configuration of chamber  130  may be tapered to expand the cooling flow at a lower pressure in the upstream direction. 
         [0040]      FIGS. 8 and 8A  illustrate yet another exemplary but nonlimiting embodiment. It will be appreciated that  FIG. 8  is a section taken transverse to the longitudinal axis of the combustor. In this view, it can be appreciated that the transfer tubes  164  may be formed as an integral part (e.g., cast or otherwise suitably formed) of a respective plurality of radially-oriented structural supports  166  that extend between the impingement sleeve assembly  168  and the transition piece  170 . The supports  166  are formed to include a radially inward inlet opening  172 , radial passageway  174  and plural exit openings  176  that permit the cooling air to flow through aligned apertures  178  in the spring fingers  180  of the hula seal  182  (only partially shown) to thereby cool the area radially inward of the hula seal  182  substantially as described above. 
         [0041]    Turning to  FIG. 9 , a simplified illustration of another cooling arrangement is provided. The combustor liner  182 , flow sleeve  184 , transition piece  186  and impingement sleeve  188  remain substantially as previously described. The aft end of the liner  182  is formed with an annular recess  190  closed on its radially outer side by an annular cover plate  192 . The plate  192  supports the annular hula seal  194  extending radially between the aft end of the plate  192  and the transition piece  186 . Each of the several transfer tubes  196  extends radially between the impingement sleeve  188  and the transition piece  186 , supplying cooling air to an area  198  behind (i.e., toward the forward end of the hula seal  194 ). This area is sealed at its forward end by a second seal  200 , forcing the cooling air to flow through the apertures  202  in the cover plate  192  and into the annular recess or chamber  190 , exiting via the apertures  204  in the cover plate  192  at the aft end of the liner and apertures  206  in the hula seal  194 . This arrangement cools the forward end of the hula seal by impingement cooling and cools the aft end of the liner by convection cooling while also purging the space  208  beneath the hula seal. The cooling air flow can be precisely controlled by optimizing the size, shape and number of transfer tubes  196 , apertures  202  and apertures  204 . 
         [0042]      FIG. 10  illustrates yet another exemplary but nonlimiting cooling arrangement. The combustor liner, flow sleeve, transition duct and impingement sleeve remain substantially as previously described. Note in this view, however, that the flow sleeve and impingement sleeve have been omitted. The aft end of the liner  210  is again formed with an annular recess  212  closed on its radially outward side by an annular cover plate  214 , with an annular hula seal  216  extending radially between the aft end of the plate  214  and the transition piece  218 . In this embodiment, the hula seal is again reversed or inverted relative to is orientation in, for example,  FIG. 9 . Cooling air from the compressor flows through the transfer tubes  220  and into the space  222  radially outward of the hula seal  216  to thereby impingement cool the seal. Cooling air then flows through apertures  224  in the spring fingers of the hula seal and through aligned apertures  226  in the cover plate, following a serpentine path into the annular recess  212 . All of the cooling air flows from the aft end of the liner toward the forward end, substantially parallel to the flow of cooling air in the aligned passages between the transition duct and impingement sleeve on the one hand, and between the combustor liner and flow sleeve on the other. The cooling air exits the recess  212  via apertures  228  at the forward end of the cover plate and joins the flow of air in the aligned passages mentioned above. It will be appreciated that the air in space  222  is purged while the hula seal is impingement cooled, and the liner aft end is cooled primarily by convection cooling. 
         [0043]      FIG. 11  illustrates yet another cooling arrangement wherein a hula seal  230  is fixed at its forward end  232  to the transition piece  234 , while an aft end  236  is resiliently compressed between the aft end of the liner  238  and the transition duct for movement relative thereto. The forward end  232  is fixed to the transition piece  234  preferably by welding, via a separate (shown) or integral (not shown) seal element  240 . In this embodiment, the seal itself serves as an impingement plate, eliminating the need for a separate cover plate as shown, for example, at  214  in  FIG. 10 . Accordingly, cooling air flowing through the transfer tube  244  will flow into the cavity  246  to cool the seal, and then flow through apertures  248  in the seal into an area  250  radially below the seal, where it impingement cools the aft end of the liner  238 . The cooling flow subsequently exits through the slot  252  at the forward end of the seal, joining the cooling air flowing in the radial passage between the flow sleeve and combustor liner to the combustors. 
         [0044]    Turning now to  FIG. 12 , an internal annular manifold  254  is formed at the aft end of the transition piece  256 , receiving the cooling air from the transfer tubes  258 . The manifold  254  supplies air through circumferentially-spaced apertures in the transition piece, and through aligned apertures  262  in the spring fingers  264  of the hula seal  266 , into the area  268  radially between the hula seal  266  and a cover plate or sleeve  270  fixed to the liner  272 . Air then flows through apertures  274  in the cover plate and exits at the forward end of the cover plate via slots  276 , joining the flow in the annular passage between the liner and the flow sleeve. 
         [0045]    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.