Abstract:
A turbine includes a transition portion where a combustor section joins a transition piece. The combustor section includes a combustor liner having an aft end that joins a transition piece body of the transition piece. A reduced thickness portion at the aft end of the combustor liner is covered by a cover sleeve to form an air flow passage on the aft end of the combustor liner. Apertures in the forward portion of the cover sleeve allow cooling air to flow into air flow passage. A plurality of turbulators project radially outward from the reduced thickness portion of the combustor sleeve towards said cover sleeve. An arch shaped resilient seal structure is positioned between the cover sleeve and the transition piece body. Supports formed on the reduced thickness portion of the combustor liner bear against the inside of the cover sleeve to prevent the cover sleeve from deforming inward due to a force applied by the seal, thereby ensuring that the air flow passage remains open.

Description:
[0001]    This application is a continuation-in-part of U.S. application Ser. No. 11/905,238 filed Sep. 28, 2007, the entire contents of which are hereby incorporated by reference. 
     
    
     BACKGROUND OF THE INVENTION 
       [0002]    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 combustion section and discharge section of the turbine. 
         [0003]    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 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. 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. 
         [0004]    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. Although lean premixed combustion is cooler than diffusion combustion, the flame temperature is still too hot for prior conventional combustor components to withstand. 
         [0005]    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 difficult at best. 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 destruction 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. 
         [0006]    With respect to the combustor liner, one current practice is to impingement cool the liner, or to provide turbulators on the exterior surface of the liner (see U.S. Pat. No. 7,010,921). 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). The various known techniques enhance heat transfer but with varying effects on thermal gradients and pressure losses. 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. Dimple concavities function by providing organized vortices that enhance flow mixing and scrub the surface to improve heat transfer. 
       BRIEF DESCRIPTION OF THE INVENTION 
       [0007]    The above discussed and other drawbacks and deficiencies are overcome or alleviated in an example embodiment by an apparatus for cooling a combustor liner and transition piece of a gas turbine. 
         [0008]    The invention may be embodied in a combustor for a turbine comprising: a combustor liner; a first flow sleeve surrounding said combustor liner with a first flow annulus therebetween, said first flow sleeve having a plurality of cooling apertures formed about a circumference thereof for directing compressor discharge air as cooling air into said first flow annulus; a transition piece body connected to said combustor liner, said transition piece body being adapted to carry hot combustion gases to the turbine; a second flow sleeve surrounding said transition piece body, said second flow sleeve having a second plurality of cooling apertures for directing compressor discharge air as cooling air into a second flow annulus between the second flow sleeve and the transition piece body, said first flow annulus connecting to said second flow annulus; a resilient seal structure disposed radially between an aft end portion of said combustor liner and a forward end portion of said transition piece body; and a cover sleeve disposed between said aft end portion of said combustor liner and said resilient seal structure, an air flow passage being defined between said cover sleeve and said aft end portion of said combustor liner, said cover sleeve having at a forward end thereof a plurality of air inlet feed holes for directing cooling air from said first annulus into said air flow passage, a radially outer surface of said combustor liner aft end portion defining said air flow passage including a plurality of turbulators projecting towards but spaced from said cover sleeve and a plurality of supports extending to and engaging said cover sleeve to space said cover sleeve from said turbulators to define said air flow passage. 
         [0009]    The invention may also be embodied in a turbine engine comprising: a combustion section; an air discharge section downstream of the combustion section; a transition region between the combustion and air discharge sections; a combustor liner defining a portion of the combustion section and transition region; a first flow sleeve surrounding said combustor liner with a first flow annulus therebetween, said first flow sleeve having a plurality of rows of cooling apertures formed about a circumference of said first flow sleeve for directing compressor discharge air as cooling air into said first flow annulus; a transition piece body connected to at least one of said combustor liner and said first flow sleeve, said transition piece body being adapted to carry hot combustion gases to a stage of the turbine corresponding to the air discharge section; a second flow sleeve surrounding said transition piece body, said second flow sleeve having a second plurality of rows of cooling apertures for directing compressor discharge air as cooling air into a second flow annulus between the second flow sleeve and the transition piece body, said first flow annulus connecting to said second flow annulus; a resilient seal structure disposed radially between an aft end portion of said combustor liner and a forward end portion of said transition piece body; and a cover sleeve disposed between said aft end portion of said combustor liner and said resilient seal structure, an air flow passage being defined between said cover sleeve and said aft end portion of said combustor liner, said cover sleeve having at a forward end thereof a plurality of air inlet feed holes for directing cooling air from said first annulus into said air flow passage, a radially outer surface of said combustor liner aft end portion defining said air flow passage including a plurality of turbulators projecting towards but spaced from said cover sleeve and a plurality of supports extending to and engaging said cover sleeve to space said cover sleeve from said turbulators to define said air flow passage. 
         [0010]    The invention may also be embodied in a method of cooling a transition region between a combustion section comprising a combustor liner and a first flow sleeve surrounding said combustor liner with a first flow annulus therebetween, said first flow sleeve having a plurality of cooling apertures formed about a circumference thereof for directing compressor discharge air as cooling air into said first flow annulus, and a transition region comprising a transition piece body connected to said combustor liner, said transition piece body being adapted to carry hot combustion gases to a turbine, a second flow sleeve surrounding said transition piece body, said second flow sleeve having a second plurality of cooling apertures for directing compressor discharge air as cooling air into a second flow annulus between the second flow sleeve and the transition piece body, said first flow annulus connecting to said second flow annulus; said transition region including a resilient seal structure disposed radially between an aft end portion of said combustor liner and a forward end portion of said transition piece body; the method comprising: configuring said aft end portion of said combustor liner so that a radially outer surface thereof includes a plurality of radially outwardly projecting turbulators and a plurality of radially outwardly projecting supports having a radial height greater than that of said turbulators; disposing a cover sleeve between said aft end portion of said combustor liner and said resilient seal structure to define an air flow passage between said cover sleeve and said aft end portion of said combustor liner, said cover sleeve having at a forward end thereof a plurality of air inlet feed holes for directing cooling air from said first annulus into said cooling air passage, said turbulators projecting towards but being spaced from said cover sleeve and said supports extending to and spacing said cover sleeve from said turbulators to define said air flow passage; and supplying compressor discharge air through at least some of said cooling apertures to and through said air inlet feed holes and through said air flow passage to reduce a temperature in a vicinity of said resilient seal. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0011]    These and other objects and advantages of this invention, will be more completely understood and appreciated by careful study of the following more detailed description of the presently preferred exemplary embodiments of the invention taken in conjunction with the accompanying drawings, in which: 
           [0012]      FIG. 1  is a partial schematic illustration of a gas turbine combustor section; 
           [0013]      FIG. 2  is a partial but more detailed perspective of a conventional combustor liner and flow sleeve joined to the transition piece; 
           [0014]      FIG. 3  is an exploded partial perspective view of the aft end of a conventional combustor liner; 
           [0015]      FIG. 4  is a cross-sectional view of the aft portion of a prior art combustor liner; 
           [0016]      FIG. 5  is a cross-sectional view of a first embodiment of the aft portion of a combustor liner having circumferential turbulators and supports; 
           [0017]      FIG. 6  is a schematic view of the aft portion of a combustor liner as illustrated in  FIG. 5 ; 
           [0018]      FIG. 7  is an enlarged cross-sectional view showing details of the encircled portion in  FIG. 5 ; and 
           [0019]      FIG. 8  is a cross-sectional view of a second embodiment of the aft portion of a combustor liner having turbulators and supports. 
       
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       [0020]      FIG. 1  schematically depicts the aft end of a combustor in cross-section. As can be seen, in this example, the transition piece  12  includes a radially inner transition piece body  14  and a radially outer transition piece impingement sleeve  16  spaced from the transition piece body  14 . Upstream thereof is the combustion liner  18  and the combustor flow sleeve  20  defined in surrounding relation thereto. The encircled region is the transition piece forward sleeve assembly  22 . 
         [0021]    Flow from the gas turbine compressor (not shown) enters into a case  24 . About 40-60% 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 compressor discharge flow passes through flow sleeve apertures  28  in the combustion liner cooling sleeve  20  and into an annulus  30  between the cooling sleeve  20  and the liner  18 . This flow of air mixes with the air from the downstream annulus  26 , and it is eventually directed into the fuel injectors inside the combustor liner  18 , where it mixes with the gas turbine fuel and is burned. 
         [0022]    In the embodiment illustrated in  FIG. 1 , the apertures  28  in the combustor flow sleeve  20  are shown as holes. In alternate embodiments, the apertures could have other shapes. For example, the apertures that admit air into the annulus  30  could be slots that extend around the circumference of the combustor flow sleeve  20 . 
         [0023]      FIG. 2  illustrates the connection at  22  between the transition piece  14 ,  16  and the combustor flow sleeve  18 ,  20 . Specifically, the impingement sleeve (or second flow sleeve) of the transition piece  14  is received in telescoping relationship in a mounting flange  32  on the aft end of the combustor flow sleeve  20  (or first flow sleeve). The transition piece  14  also receives the combustor liner  18  in a telescoping relationship. The combustor flow sleeve  20  surrounds the combustor liner  18  creating flow annulus  30  (or first flow annulus) therebetween. It can be seen from the flow arrow  34  in  FIG. 2 , that crossflow cooling air traveling in annulus  26  continues to flow into annulus  30  in a direction perpendicular to impingement cooling air flowing through the cooling apertures  28  (see flow arrow  36 ) formed about the circumference of the flow sleeve  20 . While three rows of apertures are shown in  FIG. 2 , the flow sleeve may have any number of rows of apertures. Also, as noted above, the apertures could be holes, or they could have other shapes, such as circumferential slots. 
         [0024]    Still referring to  FIGS. 1 and 2 , a typical can annular reverse-flow combustor is shown for a turbine that is driven by the combustion gases from a fuel where a flowing medium with a high energy content, i.e., the combustion gases, produces a rotary motion as a result of being deflected by rings of blading mounted on a rotor. In operation, discharge air from the compressor (compressed to a pressure on the order of about 250-400 lb/in 2 ) reverses direction as it passes over the outside of the combustor liners (one shown at  18 ) and again as it enters the combustor liner  18  en route to the turbine. Compressed air and fuel are burned in the combustion chamber, producing gases with a temperature of about 2800° F. These combustion gases flow at a high velocity into turbine section via transition piece  14 . 
         [0025]    There is a transition region indicated generally at  22  in  FIG. 1  between the combustion section and the transition piece. As previously noted, the hot gas temperature at the aft end of section  18 , the inlet portion of region  22 , is on the order of about 2800° F. However, the liner metal temperature at the downstream, outlet portion of region  22  is preferably on the order of 1400-1550° F. With reference to  FIG. 3 , to help cool the liner to this lower metal temperature range, during passage of heated gases through region  22 , the aft end  50  of the liner defines passage(s) through which cooling air is flowed. The 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. 
         [0026]    Referring to  FIG. 3 , liner  18  has an associated compression-type seal  38 , commonly referred to as a hula seal, mounted between a cover plate  40  of the liner aft end  50 , and transition piece  14 . More specifically, the cover plate  40  is mounted on the liner aft end  50  to form a mounting surface for the compression seal. As shown in  FIG. 3 , liner  18  has a plurality of axial channels  42  formed with a plurality of axial raised sections or ribs  44  all of which extend over a portion of aft end  50  of the liner  18 . The cover plate  40  and ribs together define the respective airflow channels  42 . These channels are parallel channels extending over a portion of the aft end of liner  18 . Cooling air is introduced into the channels through air inlet slots or openings  46  at the forward end of the channels. The air then flows into and through the channels  42  and exits the liner through openings  48 . Alternatively, or in addition, cooling air may enter the channels  42  through apertures or holes  47  in the cover plate  40 . As shown in  FIG. 4 , the cross-section of the channel as defined by its height may decrease along the length of the channel in an aft direction. 
         [0027]    As noted, the invention pertains to the design of a combustor liner used in a gas turbine engine and more specifically the cooled aft-end of the combustor liner as an improvement to the conventional structure shown in  FIG. 4 . As noted above, this area has conventionally been composed of axial grooves  42  machined into the liner  18  and a sheet metal cover  40  to support the aft-end Hula seal  38 . 
         [0028]    According to an example embodiment of the invention, rather than providing axial grooves  42  as in the conventional combustor liner, an annular cooling system is provided that features transverse turbulators  142  as illustrated in  FIGS. 5-7 . As illustrated in  FIG. 5 , a sheet metal cover  140  is provided to support the aft-end Hula seal  38 . The cover  140  defines an air passage with the liner aft-end  150 . The sheet metal cover  140  includes air inlet apertures  146  for passage of cooling media to the region below the Hula seal  38 . Spaced supports  144  are provided on the aft-end of the combustor liner  150  under the forward and aft ends of the Hula seal  38  to keep the sheet metal cover  140  spaced from the liner aft-end  150 . 
         [0029]    As illustrated in  FIG. 6 , although the supports  144  extend around the circumference of the liner  150 , gaps  143  are formed between the individual supports  144 , the gaps  143  being circumferentially spaced from one another about the axis of the combustor liner. In the illustrated embodiment, four axially spaced rows of supports  144  are provided, as shown in  FIG. 5 , each row comprised of a plurality of circumferentially spaced supports  144 , as shown in  FIG. 6 . 
         [0030]    Advantages of the illustrated design are many in comparison with the conventional design of  FIG. 4  and include better heat transfer per unit air used, easier production than axial grooves from a machine/manufacturing standpoint; lower heat input to the temperature limited Hula seal; and an ability to achieve a lower temperature in the liner&#39;s aft end, which would be critical in engines with higher firing temperatures. 
         [0031]    The transverse turbulators  142  provided according to an example embodiment of the invention are a highly effective heat transfer augmentation device. It is common to see heat transfer numbers of about 200% better than non-turbulated sections with the same quantity of cooling air. Therefore, by providing transverse turbulators  142  as proposed herein, it is possible to achieve the same amount of heat transfer as in the conventional structure with less cooling air. This would be a highly desirable feature in lean pre-mixed gas turbines because the cooling air can be used more effectively in other parts of the system. The transverse turbulators are expected to be more manufacturing friendly than the conventional axial channels because, in particular they are less sensitive to small variations in the manufacturing process then channeled flow. 
         [0032]    As noted above, among current cooing systems are those composed of numerous axially extending cooling channels. These channels  42  are defined by walls that extend radially outward from the cold side of the liner aft end  50  to the sheet metal cover  40 , as shown in  FIG. 4 . The cover  40  makes contact with and is supported by the top of the channel defining walls  44  (see U.S. Pat. No. 7,010,921). A significant amount of heat transfer flows through this assembly and into the Hula seal  38  that sits on top of the sheet metal cover  40 . 
         [0033]    The Hula seal&#39;s function is to act like a spring while maintaining a good seal. This part has a limited temperature capability and is often very close to its functional limit. The configuration proposed herein ( FIGS. 5-7 ) helps limit the amount of heat transferred to the Hula seal by significantly reducing the contact area through which the heat can flow into the seal by limiting that contact area to the spaced supports  144 . 
         [0034]    An alternate embodiment is illustrated in  FIG. 8 . In this embodiment, the Hula seal  38  is rotated 180° from the position it occupied in the embodiment illustrated in  FIGS. 5-7 . As a result, only the center arched portion of the seal  38  bears against the top of the cover  140 . The ends of the Hula seal  38  would then bear against the forward end of the inner sleeve  14  of the transition piece  12 . 
         [0035]    This embodiment only requires two circumferential rows of supports  144  located under the arched center portion of the Hula seal  38 . In still other embodiments, only a single circumferential row of supports may be provided under the arched center portion of the Hula seal  38 . Because an embodiment as illustrated in  FIG. 8  requires fewer circumferential rows of supports  144 , the cost and time required to manufacture the combustor liner  150  can be reduced compared to the embodiment illustrated in  FIGS. 5-7 . 
         [0036]    In addition, in this embodiment only one or two rows of the supports  144  would act to transfer heat from the combustor liner  150  to the cover plate  140 , and then into the Hula seal. Thus, the embodiment illustrated in  FIG. 8  provides even less of a pathway for heat to be transferred to the Hula seal  38 , which should further serve to keep the Hula seal at a desirably low temperature. 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.