Abstract:
A gas turbine combustor structure having improved cooling effectiveness and increased life as well as a method for improving the cooling effectiveness is disclosed. The gas turbine combustor incorporates a unique flow sleeve configuration for directing air to more effectively cool a combustion liner. The flow sleeve geometry is configured to incorporate a conical aft portion having a plurality of air feed holes that reduce pressure loss to the incoming air and flow separation effects from the surrounding combustor hardware, thereby resulting in improved combustor performance.

Description:
BACKGROUND OF THE INVENTION 
   1. Field of the Invention 
   The present invention relates to gas turbine combustors and more specifically to a flow sleeve having an inlet region that reduces pressure loss to the compressed air entering a combustor. 
   2. Description of Related Art 
   A gas turbine engine typically comprises a multi-stage compressor, which compresses air drawn into the engine to a higher pressure and temperature. A majority of this air passes to the combustors, which mixes the compressed heated air with fuel and contains the resulting reaction that generates the hot combustion gases. These gases then pass through a multi-stage turbine, which drives the compressor, before exiting the engine. In land-based gas turbines, the turbine is also coupled to a generator for generating electricity. 
   For land-based gas turbine engines, often times a plurality of combustors are utilized. Each of the combustion systems include a case that serves as a pressure vessel containing the combustion liner, which is where the high pressure air and gas mix and react to form the hot combustion gases. Typically the case is fabricated from a lower temperature capable material such as carbon-steel. In order to ensure that the case is not overexposed to the temperatures of the combustion liner as well to ensure that the combustion liner receives the proper amount of air for cooling and mixing with the fuel, an additional liner is often located within the case and is coaxial to the combustion liner and case. This additional liner is more commonly referred to as a flow sleeve. 
   A two-stage combustion system of the prior art commonly used in land-based gas turbine engines is shown in cross section in  FIG. 1 . Combustor  10  includes a generally annular case  11  having a center axis A—A and an end cover  12  that is fixed to a case flange and contains a plurality of fuel nozzles  13  located about center axis A—A. Located coaxial to center axis A—A is a combustion liner  14  having a first combustion chamber  15  and second combustion chamber  16 , separated by venturi  17  having a throat of reduced cross sectional area  18 . An additional fuel nozzle  19  is located along center axis A—A. Located coaxial to combustion liner  14  and radially between case  11  and combustion liner  14  is flow sleeve  20 . As mentioned previously, flow sleeve  20  serves to direct compressed air along the outer walls of liner  14  for cooling purposes, as well as for being injected to mix with the fuel for combustion. In combustor  10  of the prior art, flow sleeve  20  forms a generally annular passageway  21  around combustion liner  14  for directing the required amount of compressed air to combustion liner  14  for cooling and mixing with the fuel from fuel nozzles  13  and  19 . In prior art combustor  10 , compressed air is introduced to the combustion system through a generally annular flow sleeve inlet  22 , which is shown in a more detailed cross section in  FIG. 2 . 
   Flow sleeve inlet  22  is formed between flow sleeve  20  and transition duct  25 , which has a bellmouth portion  26  and a structural support ring  27 , each of which are located towards the forward end of transition duct  25 . In this combustor configuration, bellmouth  26  and support ring  27  create obstructions that block or disturb a portion of the compressed air flow that enters passageway  21  through flow sleeve inlet  22 , thereby causing an undesirable pressure loss to the air supply. This disturbance to the air flow and resulting pressure loss has multiple negative effects on the hardware durability and performance. Specifically, hula seal  28 , which, in the prior art, is a seal encompassing the aft end outer surface of liner  14  and contains a plurality of axial slots that form “fingers” that spring to seal between liner  14  and transition duct  25 , does not receive sufficient cooling air due to a separation zone  29  created by air flow passing over bellmouth  26  (see  FIG. 2 ). As a result of this lack of cooling air, the aft end of combustion liner  14  and hula seal  28  operate at a higher temperature, causing more radial interference between hula seal  28  and transition duct  25  than desired, leading to premature wear of hula seal  28 . The flow disturbances created by bellmouth  26  and ring  27  combined with the geometry of flow sleeve inlet  22 , due to the axial length of the aft region of flow sleeve  20 , creates a pressure loss to the incoming air supply. The pressure loss at flow sleeve inlet  22 , which is approximately 1.5% of the available air pressure, results in a lower cooling air supply pressure to combustion liner  14 . Annular passageway  21  creates little, if any, additional pressure loss to the cooling air. As a result, less air is passed through the various passages requiring cooling and injected for mixing with the fuel, thereby resulting in higher operating temperatures, a less durable design, and reduced combustor performance. As one skilled in the art of gas turbine combustion will understand, maintaining adequate cooling of the combustion liner is imperative for combustor durability and performance. 
   Therefore, what is needed is a flow sleeve for a gas turbine combustor having an inlet region that reduces the pressure loss to the incoming compressed air, such that a high enough air pressure is available to provide sufficient cooling to the combustion liner surfaces. This is especially true for combustors that operate for an extended period of time and require large amounts of cooling and enhanced mixing in order to achieve low emissions. 
   SUMMARY AND OBJECTS OF THE INVENTION 
   A gas turbine combustor structure having improved cooling effectiveness and increased life as well as a method for improving the cooling effectiveness is disclosed. The gas turbine combustor in accordance with the preferred embodiment of the present invention comprises a generally cylindrical case that serves as a pressure vessel having a generally cylindrical end cover fixed to a first case flange. The end cover has a plurality of first fuel nozzles arranged about a center axis. Located within the case and coaxial to the center axis is a flow sleeve that is used to direct compressed air along a combustion liner for cooling and injection into the liner. The flow sleeve has a first portion that is generally cylindrical in shape, a mounting flange for mounting the flow sleeve to a second case flange, and a second portion that is generally conical in shape that is fixed to the first portion of the flow sleeve. The second portion of the flow sleeve contains a plurality of feed holes for supplying cooling air to a generally annular passageway that is formed between the flow sleeve and the combustion liner. The combustion liner is in fluid communication with a plurality of fuel nozzles and is supplied with air from the generally annular passageway for cooling of the liner walls as well as for mixing with fuel that is injected from the fuel nozzles. Hot combustion gases formed in the combustion liner are directed towards the turbine section by way of a transition duct. In order to prevent hot gases from leaking, the combustion liner seals to the transition duct by a seal located proximate the liner aft end outer wall that has a means for passing cooling air through the seal to cool beneath the seal. 
   The present invention avoids the shortcomings of the prior art by providing an improved flow sleeve design that reduces the pressure loss to the cooling air at the flow sleeve inlet, by approximately 50%, thereby providing the combustion liner with higher pressure air for cooling and mixing with fuel for combustion. This is accomplished by altering the flow sleeve inlet region such that all air enters the flow sleeve upstream of the transition duct and a majority of that air enters the flow sleeve through a plurality of feed holes in the conical portion of the flow sleeve. Moving the air inlet location away from the transition piece bellmouth and support ring as well as reconfiguring the inlet geometry, eliminates a majority of the pressure losses associated with the prior art configuration. 
   It is an object of the present invention is to provide a gas turbine combustor having lower pressure losses to the cooling air supply pressure. 
   It is another object of the present invention to provide a method of improving the cooling effectiveness of an aft region of a combustion liner. 
   It is yet another object of the present invention to provide a gas turbine combustor having improved durability as a result of the lower pressure losses to the cooling air supply. 
   In accordance with these and other objects, which will become apparent hereinafter, the instant invention will now be described with particular reference to the accompanying drawings. 

   
     BRIEF DESCRIPTION OF DRAWINGS 
       FIG. 1  is a cross section view of a gas turbine combustor of the prior art. 
       FIG. 2  is a detailed cross section view of the flow sleeve inlet region of a gas turbine combustor of the prior art. 
       FIG. 3  is a cross section view of a gas turbine combustor in accordance with the preferred embodiment of the present invention. 
       FIG. 4  is a detailed cross section view of the flow sleeve inlet region of a gas turbine combustor in accordance with the preferred embodiment of the present invention. 
       FIGS. 5A and 5B  are elevation views of a portion of the aft section of a combustion liner and seal, including a means for passing cooling air through the seal, in accordance with the preferred embodiment of the present invention. 
   

   DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT 
   The preferred embodiment of the present invention is shown in detail in  FIGS. 3–5B . Gas turbine combustor  40 , in accordance with the present invention comprises a generally cylindrical case  41  having center axis B—B, first case flange  42 , and second case flange  43 . Fixed to first case flange  42  is a generally cylindrical end cover  44  that has a plurality of first fuel nozzles  45  arranged in an annular array about center axis B—B. Located radially within case  41  and coaxial to center axis B—B is flow sleeve  46  having first portion  47 , second portion  48 , and mounting flange  49 . First portion  47  is generally cylindrical in shape and has a first end  50  located proximate first case flange  42 . Mounting flange  49  extends radially outward from first portion  47  and is located axially along first portion  47  proximate second case flange  43 , and fixes flow sleeve  46  to case  41  at second case flange  43 . For the preferred embodiment, second end  51  of first portion  47  is located proximate mounting flange  49 . Flow sleeve  46  also includes second portion  48 , which is generally conical in shape, and has a first end  52 , which is fixed to second end  51  of first portion  47 , and a second end  53  having an inlet ring  54 . Located around the perimeter of second portion  48  is a plurality of feed holes  55 . The location and size of the feed holes can vary depending on the required air flow, but for the preferred embodiment, the feed holes are arranged in at least one row about second portion  48  of flow sleeve  46 . 
   Located within flow sleeve  46  and coaxial to center axis B—B is a generally annular combustion liner  56  that is in fluid communication with first fuel nozzles  45  and a second fuel nozzle  45 A. Combustion liner  56  comprises an inner wall  57 , an outer wall  58 , a first liner end  59 , and a second liner end  60 , with a seal  61  fixed to and encompassing outer wall  58  proximate second liner end  60 . Seal  61 , which seals against transition duct  62 , also includes a means for passing cooling air through seal  61 . The sealing interface region and aft end of combustion liner  56  is shown in greater detail in  FIG. 4 . Further details regarding the means disclosed for passing cooling air through seal  61  is shown in  FIGS. 5A and 5B . Specifically, two configurations are shown that each comprise a plurality of openings  63  that pass a first supply of cooling air through seal  61  to cool outer wall  58  of combustion liner  56  proximate second liner end  60 . In order to provide surface cooling to inner wall  57  proximate second liner end  60 , a second supply of cooling air is directed along inner wall  57  for cooling the aft end region of combustion liner  56 . The second supply of cooling air can be directed along inner wall  57  by a variety of means, most commonly through a plurality of precisely sized cooling holes located in combustion liner  56  proximate the region requiring cooling. 
   The cooling air (CA) entering the flow sleeve inlet region is used for three purposes proximate the aft end of combustor  40 . Each of these locations benefit from the flow sleeve redesign to reduce the pressure loss to the cooling air. Referring now specifically to  FIG. 4 , aft end of combustor  40  is shown in detail and includes a plurality of arrows indicating the cooling air (CA) and its various directions. A first supply of cooling air, CA 1 , is directed between bellmouth  65  of transition duct  62  and inlet ring  54  of flow sleeve  46 . First supply of cooling air CA 1  is directed through plurality of openings  63  in seal  61  to cool outer wall  58  of combustion liner  56  in the region beneath seal  61  and area proximate second liner end  60 . The quantity and configuration of openings  63  in seal  61  depends on the amount of air required in order to achieve sufficient cooling. As shown in  FIG. 5 , openings  63  can take on different configurations, such as holes or slots. 
   A second supply of cooling air CA 2  is primarily directed through feed holes  55  in second portion  48  of flow sleeve  46  and is injected into combustion liner  56  at a region requiring cooling along inner wall  57 . The exact location and orientation of the injected air depends on the combustion liner operating conditions and amount of available cooling air. The location of feed holes  55  ensures a sufficient supply of cooling air with minimal pressure loss since feed holes  55  are placed upstream of transition duct bellmouth  65  and support ring  66 , such that any flow disturbance from the bellmouth or support ring are insignificant. 
   A third supply of cooling air CA 3  is directed through feed holes  55  in second portion  48  of flow sleeve  46  and along outer wall  58  and towards first liner end  59  for cooling combustion liner  56  and for mixing with fuel from fuel nozzles  45  inside combustion liner  56 . Feed holes  55  are sized such that the pressure drop across the feed holes is minimized, thereby supplying a higher air pressure to the cooling and combustion process than the prior art gas turbine combustor. This is especially imperative when cooling a dual stage combustor that incorporates an effusion cooled combustion liner and a counter flow venturi, similar to that shown in  FIG. 3 , and disclosed in U.S. Pat. Nos. 6,427,446, 6,446,438, and 6,484,509, assigned to the same assignee herein. In this type of combustion system, cooling air is drawn in to venturi cooling passageway  70  proximate venturi aft end  71  and is injected into a chamber  72  upstream of the venturi throat  73  for mixing with the fuel and air, such that the fuel/air mixture is leaner, resulting in lower emissions. When cooling a venturi in this manner, the temperature of the cooling air rises dramatically while the air pressure drops as it passes through venturi cooling passageway  70 , prior to being injected into chamber  72 . Flow throughout venturi cooling passageway  70  relies on pressure changes to pass the cooling air from venturi aft end  71  to chamber  72 . Therefore, given the known pressure losses to occur in this system, the air entering venturi cooling passageway  70  must initially have a higher pressure in order to adequately cool the venturi system and be injected into chamber  72  for mixing with fuel for combustion. This higher air pressure is possible due to the redesigned second portion geometry that moves the air inlet region forward of the transition duct bellmouth  65  and support ring  66 , such that the inlet region is removed from any disturbances created by either of these structures while also introducing a majority of the air through a plurality of feed holes  55 . 
   Inherent in the aforementioned gas turbine combustor structure is a method of improving the cooling effectiveness and increasing component life of a combustion liner aft region. The method comprises the steps of providing a gas turbine combustor  40  having a case  41  with first case flange  42  and second case flange  43 , a transition duct  62 , a flow sleeve  46  with a first portion  47  generally cylindrical in shape, having a first end  50 , a second end  51 , and a mounting flange  49  for securing flow sleeve  46  to second case flange  43 , and a second portion  48  generally conical in shape having a first end  52 , a second end  53 , and a plurality of feed holes  55 . First end  52  of said second portion  48  is fixed to second end  51  of said first portion  47  and second end  53  of second portion  48  has an inlet ring  54 . Gas turbine combustor  40  also has a combustion liner  56 , that is located radially within flow sleeve  46 , and has an inner wall  57 , an outer wall  58 , a first liner end  59 , a second liner end  60 , and a seal  61 , having a means for passing cooling air through seal  61 , fixed to outer wall  58  proximate second liner end  60 . Preferably, means for passing cooling air through seal  61  comprises a plurality of openings  63 , which can be a variety of configurations, including holes or slots. 
   Next, a first supply of cooling air, CA 1 , passes through an opening between flow sleeve support ring  54  transition duct  62  and is directed through plurality of openings  63  in seal  61  to cool outside wall  58  of combustion liner  56  and the region beneath seal  61 . Also, a second supply of cooling air, CA 2 , which passes primarily through plurality of feed holes  55 , is injected into combustion liner  56  and directed along inner wall  57  for cooling purposes. Typically cooling air CA 2  enters combustion liner  56  through a plurality of cooling holes whose location depends on the combustor configuration. Finally, a third supply of cooling air, CA 3 , which also passes through plurality of feed holes  55 , is directed along outer wall  58  of combustion liner  56  for additional liner aft end cooling as it flows towards venturi cooling passageway  70  and first liner end  59 . Each of the cooling air supplies CA 1 , CA 2 , and CA 3  are supplied to combustor  40  at a higher pressure than in prior art combustors due to the redesigned flow sleeve second portion  48 , including feed holes  55 , and its location relative to transition duct  62 . 
   While the invention has been described in what is known as presently the 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 within the scope of the following claims.