Abstract:
An improved dual-stage, dual-mode turbine combustor capable of reducing nitric oxide (NOx) emissions is disclosed. This can-annular combustor utilizes multiple, single wall sheet metal combustor liners, generally annular in shape, and each liner having multiple hole film cooling means, which includes at least one pattern of small, closely spaced film cooling holes angled sharply in the downstream direction and various circumferential angles for improved liner cooling and improved fuel/air mixing within the liner, resulting in lower NOx emissions.

Description:
BACKGROUND OF THE INVENTION 
     1. Field of the Invention 
     This invention relates generally to an apparatus and method for reducing nitric oxide (NOx) emissions and cooling the combustion liner for a can-annular gas turbine combustion system. Specifically, an apparatus and method for introducing the cooling air into the premix chamber of the combustion system that minimizes the use of compressor discharge air for cooling the combustion liner as well as for improving the mixing of fuel and air prior to the combustion process. 
     2. Description of Related Art 
     Combustion liners are commonly used in the combustion section for most gas turbine engines. The combustion section is located between the compressor and turbine, and depending upon the application, the combustion section may not be located along the centerline of the engine, but may be located around the centerline or even perpendicular to the engine orientation. Combustor liners serve to protect the combustor casing and surrounding engine from the extremely high operating temperatures by containing the chemical reaction that occurs between the fuel and air. 
     Recent government emission regulations have become of great concern to both gas turbine manufacturers and operators. Of specific concern is nitric oxide (NOx) due to its contribution to air pollution. While NOx emissions are of some concern to aircraft engines, greater concerns include engine weight, performance, safety and fuel efficiency. While these concerns are shared by the industrial gas turbine engine market, NOx emissions rank as one of the greatest concerns. Utility sites have governmental permits that allow specific amounts of emissions each year. Lower emission rates, especially NOx, allow engines to run longer hours and hence generate more revenue. 
     It is well known that NOx formation is a function of flame temperature, air inlet temperature, residence time, and equivalence ratio. Nitric oxide emissions have been found to be lower for lower flame temperature, lower air inlet temperature, shorter residence time, and lower equivalence ratio, or a leaner fuel mixture. Lower flame temperatures and lower equivalence ratios can be accomplished by increasing the amount of air used in the combustion process, for a given amount of fuel. Further reductions in emissions can be accomplished by improving the utilization of the cooling air. 
     The present invention is used in a dry, low NOx gas turbine engine, which is typically used to drive electrical generators. Each combustor includes an upstream premix fuel/air chamber and a downstream, or secondary, combustion chamber, separated by a venturi having a narrow throat constriction. The present invention is concerned with improving the mixing of fuel and air in the premix zone as well as the cooling of the combustion liner to further reduce nitric oxide emissions. 
     Prior methods of cooling combustion liners vary extensively. U.S. Pat. No. 4,292,801 and U.S. Pat. No. 5,127,221 describe louver film cooling and transpiration cooling, respectively, for similar dual-stage, dual-mode combustors. Backside impingement cooling is described in U.S. Pat. No. 5,117,636. Though these methods of cooling have proven adequate throughout the engine operating cycle, enhancements can be made to further reduce pollutants from the combustor, while improving cooling effectiveness. 
     Over the years, some annular gas turbine combustor designers have incorporated angled film cooling holes, specifically for improving cooling efficiency. Typically, annular combustors are used for aircraft applications where small size and reduced weight are important design factors. Angled film cooling holes improve cooling efficiency by increasing the amount of internal surface area that is available for heat removal. For example, a hole drilled at 20 degrees to the liner wall has nearly three times as much surface area as a hole drilled normal to the liner surface. In addition, angled film cooling holes provide a jet of air to form a better film along the liner surface. In order to accomplish this improved cooling, thicker liner walls are typically required, which further increase hole surface area, hence an increase in liner weight. Examples of annular aircraft combustors utilizing this cooling technique are discussed in U.S. Pat. No. 5,233,828; U.S. Pat. No. 5,181,379; U.S. Pat. No. 5,279,127; and U.S. Pat. No. 5,261,223. This technique is also used in an annular liner dome plate as described in U.S. Pat. No. 5,307,637, and to provide differential cooling to accommodate hot spots on annular combustor liner surfaces, as discussed in U.S. Pat. No. 5,241,827. 
     Of greater importance to reduce NOx emission than the improved cooling is the improved mixing of the air with fuel for combustion. When cooling performance is improved, less air is typically required for cooling and more can be dedicated to fuel/air mixing. More air into the combustion process will lower fuel to air ratios and hence equivalence ratio as well as lower flame temperature, which, as explained earlier, are two key drivers of NOx emissions. The increased air for the combustion process can be delivered through the front end of the combustor with the fuel or through the cooling holes as part of the jet. The jet of air would then provide the cooling film for the liner surface as well as a jet of air to mix with the fuel prior to combustion. This increase in mixing performance can be improved further by angling the holes in the circumferential direction to induce a swirl within the combustor. 
     The present invention provides for improved combustor cooling while enhancing fuel/air mixture in the combustor for a dual-stage, dual-mode low NOx combustor with a dedicated premix chamber. 
     BRIEF SUMMARY OF THE INVENTION 
     An improved apparatus and method for mixing fuel and air, while at the same time cooling a gas turbine combustion liner in a can-annular low NOx gas turbine engine that includes a gas turbine combustor having a premixing chamber, a secondary combustion chamber with a venturi, described as a dual-mode, dual-stage combustor. Each gas turbine engine typically has a plurality of combustors. 
     In accordance with the present invention, each can-annular combustion liner is substantially cylindrical and includes an array of multiple film cooling apertures and dilution cooling apertures disposed in a predetermined array and direction of air flow, resulting in improved cooling performance on the combustion liner, while at the same time providing improved fuel and air mixture in the combustor. 
     The array of multiple film holes in each can-annular combustion liner includes angling each of the film cooling holes or apertures, both in an axial direction and a circumferential direction. The directionality of air flowing through the angled holes provides for a predetermined flow pattern within the combustion liner that aids in fuel/air mixing. The combustion liner apertures and holes are produced by drilling holes through the combustion liner at a predetermined angular slant in the direction of combustion flow, cold side to hot side. The predetermined strategic slanted or angled aperture is not perpendicular to the combustor wall. A predetermined angle that is in two directions, both axially and circumferentially, is selected to increase the amount of surface area of the combustion liner that is being cooled, while at the same time providing directionality of flow that greatly enhances the mixing of fuel and air. The slanted holes are drilled at a circumferential angle that is preferably in the direction of combustor swirl from the premix chamber. The diameter of each of the holes and the spacing of the holes from each other is sized to maximize the cooling effectiveness of the hole pattern, improve fuel/air mixing, while at the same time not sacrificing the structural integrity of the combustion liner. 
     The apparatus described in this invention may include the combustor venturi section and air cooling flow as described in U.S. patent application Ser. No. 09/605,765 entitled “Combustion Chamber/Venturi Cooling For A Low NOx Emission Combustor” assigned to the same assignee as the present invention. The combustion liner also contains a dome section, which engages the fuel nozzles and provides another means for introducing air into the combustion process. 
     It is an object of the present invention to reduce the nitric oxide (NOx) emissions in a gas turbine combustion system by improving fuel/air mixing and lowering flame temperature. 
     It is another object of the present invention to provide a can-annular low emissions combustor system having combustion liners with apertures or holes for cooling and fuel/air mixing that are slanted axially and circumferentially. 
     It is yet another object of the present invention to incorporate an improved venturi section that utilizes its cooling air for the combustion process, further reducing polluting emissions. 
     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 THE SEVERAL VIEWS OF THE DRAWINGS 
     FIG. 1 shows a side elevation view in cross section of a typical gas turbine engine. 
     FIG. 2 shows a side elevation view in cross section of a partial gas turbine engine combustion system that represents the prior art, which utilizes louver cooling. 
     FIG. 3 shows a side elevation view in cross section of a partial gas turbine engine combustion system that represents the prior art, which utilizes transpiration cooling. 
     FIG. 4 shows a side elevation view in cross section of a partial gas turbine engine combustion system that represents the prior art, which utilizes impingement cooling. 
     FIG. 5 shows a perspective view in partial cross section of an annular aircraft gas turbine combustion system that represents prior art, which utilizes film cooling. 
     FIG. 6 shows a perspective view in partial cross section of an annular aircraft gas combustor that represents prior art, which utilizes film cooling. 
     FIG. 7 shows a gas turbine combustion liner in perspective view in accordance with the present invention. 
     FIG. 8 shows a side elevation view in cross section of a partial gas turbine combustion liner in accordance with the present invention. 
     FIG. 9 shows greater detail of a side elevation view in cross section of a partial gas turbine combustion liner in accordance with the present invention. 
     FIG. 10 shows a perspective view, partially cut away, of the premix chamber wall having angled film cooling holes in accordance with the present invention. 
    
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT 
     Referring to FIG. 1, an existing gas turbine engine  10  is shown. The engine is comprised of an air inlet  11 , a multi-stage axial compressor  12 , can-annular combustor  13 , which surrounds the aft end of the compressor, combustion transition pieces  14 , which direct combustion discharge gases into the turbine section, a multi-stage axial turbine  15  and exhaust plenum  16 . The turbine  15 , which drives compressor  12 , is connected to the compressor through an axial drive shaft  17 . This drive shaft is also coupled to the generator, which is not shown. The gas turbine engine  10 , which is primarily used for generating electricity draws air into the system through inlet  11  and is then fed into compressor  12  where it passes through multiple stages of fixed and rotating blades. The air, which is now at a much higher pressure is directed into combustion section  13 , where fuel is added and mixed with the air to form the hot gases necessary to turn turbine  15 . The hot gases exit the turbine through multiple transition pieces  14 , which direct the flow into turbine  15  at the proper orientation. The hot gases then pass through multiple stages of fixed and rotating airfoils in turbine  15 , which may or may not be cooled by bleed air drawn off of compressor  12 . The hot gases are then directed from the turbine  15  to exhaust plenum  16 . 
     Referring now to FIG. 2, a portion of a gas turbine, dual stage combustion chamber  30  is shown in cross section. The combustion liner  31  is shown inside case  36  with cover  35  installed on case  36 . Cover  35  includes multiple fuel nozzles  34  arranged in a circular pattern around the cover as well as a central fuel nozzle of similar configuration. The combustion liner  31  is a dual-stage combustor comprising a premix chamber  38  and a secondary combustion chamber  39 . The two chambers are separated by a venturi  32  with a throat  33  for the purpose of maintaining the flame in a secondary combustion chamber  39 . In this example of prior art, the liner is cooled by air passing through cooling holes  37  and directed downstream by louvers  40 . Louvers provide a rigid surface that results in increased liner structural integrity, while providing a means of directing the cooling air downstream. Louver cooling, also known as rolled ring or splash cooling, bleeds air through small rows of holes  37  in liner wall  31 , and directs it along the liner wall surface by means of an internal deflector, or louver  40 . Drawbacks to this configuration include steep temperature gradients between the metal surrounding the cooling air and the louver edge because the air from the previous row of cooling holes has lost its effectiveness. This thermal problem can produce high stresses in the liner shell resulting in cracking and extreme coating degradation. 
     A similar combustor, and another form of prior art, is shown in FIG.  3 . Again, a combustion liner  51  is shown for a dual-stage combustor comprising a premix chamber  58  and a secondary combustion chamber  59 . The two chambers are separated by a venturi  52 . All other features are similar to those described in FIG. 2, with the exception of the cooling method for venturi  52 . Louver cooling is utilized in premix chamber  58  by way of cooling apertures  57  and deflectors  61 . The venturi  52  is cooled by transpiration cooling instead of louver cooling as shown in FIG.  2 . In transpiration cooling, air or another cooling fluid passes through a porous structure, such as the venturi walls  63  into the venturi boundary layer on the hot gas path  62 . This allows the venturi inner wall  63  metal temperature to be maintained under that of the gas path  62 . Cooling of venturi  52  is accomplished by the absorption of heat from venturi walls  63  and by altering the boundary layer along venturi flowpath  62 . In order to provide adequate transpiration cooling, material for the venturi walls is typically composed of a porous metal laminate such as Lamilloy. A major drawback to this cooling method is the availability of porous materials required to provide adequate heat transfer and extended durability of these materials. 
     Referring now to FIG. 4, a similar dual-stage combustion chamber that utilizes impingement cooling is shown. Again, a combustion liner  81  is shown for a dual-stage combustor comprising a premix chamber  86  and a secondary combustion chamber  87 . The two chambers are separated by a venturi  82  with a narrow throat  85 . Other features are similar to those described and detailed in FIGS. 2 and 3. The focus of this example of prior art is the cooling method of venturi  82 . The premix section is cooled with louvers (not shown) as described in FIG.  2 . Venturi  82  is cooled by impingement of air from outside the liner shell along the backside of the venturi flowpath walls  91  and  92 . Cooling air enters the venturi section through apertures  83  in liner  81 . The air then passes through multiple rows of holes  88  and impinges on the backside of venturi gas path walls  91  and  92 . The cooling air then travels downstream through channel  93  and enters the dilution zone  90  of combustion liner  81  as film cooling. Though this cooling method has proven adequate, the major drawback to this configuration is the requirement for double wall construction to create the impingement jets, hence an increased cost and weight, as well as the extreme temperature differences created between the two venturi walls resulting in differential thermal expansion that can lead to buckling. In addition, combustion efficiency is somewhat lower due to the cooling being discharged into the dilution zone aft of the combustion region. 
     FIG. 5 shows an annular aircraft combustor  100  that utilizes angled film cooling holes for the purpose of improving cooling effectiveness of liner  101 . Multiple fuel nozzles  102  are incorporated in combustion liner  101 . Multiple rows of angled cooling holes  103  are located on the inner and outer liner skin. In addition, larger dilution holes  104  are located further downstream in the liner. A close-up view in cross section of this liner surface is shown in FIG.  6 . For the purpose of more effective liner cooling, an array of cooling holes  103  are drilled a diameter D at an axial angle A relative to the liner skin. The resulting hole is length L. Drilling holes at an angle relative to the flow path provides increased internal surface area for heat removal as well as providing a better layer of film cooling. Holes are spaced a predetermined distance S apart. The holes are also drilled at a tangential angle B to induce a swirl. 
     The present invention is disclosed in FIGS. 7,  8 ,  9 , and  10 . The combustion liner assembly  200  is a dual-stage, dual-mode low nitric oxide (NOx) combustor composed of outer liner  201 , dome cap assembly  204 , and a venturi, which is not visible in FIG.  7 . Liner  201  is held in the combustion system by forward locating tabs  203  and an aft spring seal  202 . The dome cap assembly  204  is held in outer liner  201  by pins  205 . The venturi (not shown in FIG. 7) is held in place in outer liner  201  by two rows of pins  206 . The liner shell  201  has a number of apertures at the forward end for cooling the liner wall and premixing of fuel and air for combustion. The dome cap assembly and venturi are shown in greater detail in FIGS. 8 and 9. 
     FIG. 8 shows a partial cross-section of the present invention. Again, liner shell  201  is shown with dome cap assembly  204  installed by pins  205  and venturi  212  installed via pins  206 . The dual-stage combustor is comprised of premix chamber  211  and secondary combustion chamber  210 . The dome cap assembly&#39;s primary features include openings  213  for multiple fuel nozzles located around the combustor centerline, with an additional opening  214  along the combustor centerline for a secondary fuel nozzle. This center opening  214  includes a swirler  215 . The multiple fuel nozzle receptacles  213  engage a formed dome  216 , which serves as a regulator for controlling the amount of air that enters the combustor. Venturi  212  is a separate component formed of numerous sheet metal pieces with the purpose of forming the secondary combustion chamber  210  and a narrow constriction or throat  219  that maintains the flame in secondary combustion chamber  210 . The venturi has a built-in cooling channel  220  that is formed by two cylindrical inner and outer walls,  221  and  222 , respectively, as well as a forward end  250  and an aft end  251 . The venturi, its cooling circuit, and basic operation are discussed in detail in U.S. patent application Ser. No. 09/605,765, filed Jun. 28, 2000, entitled “Combustion Chamber/Venturi Cooling for a Low NOx Emission Combustor,” assigned to the same assignee as the present invention and incorporated herein by reference. 
     The basic cooling air flow path is shown in the lower half of FIG. 8, where cooling air flow direction is represented by arrows. The cooling air travels towards the forward end of combustion liner  201 . A predetermined amount of cooling air enters cooling channel  220  through holes  223  in liner  201 . Cooling air travels upstream through channel  220  to the leading edge of venturi  212  and exits the venturi through matched holes  224  in the venturi outer skin and liner  201 . The cooling air, which has been preheated as a result of cooling the venturi inner walls  221 ,  228 , and  229 , enters an annular cavity  226  formed by a belly band  225  around liner  201 . Due to the pressure loss along cooling channel  220 , additional cooling air is supplied to annular cavity  226  by resupply holes  230 . The cooling air, now at a higher pressure is directed out of annular cavity  226  through multiple rows of angled holes  227 . This air is then premixed with the fuel and air in premix chamber  211  and used in the combustion process in secondary combustion chamber  210 . 
     The remainder of the cooling air that is not utilized in cooling venturi  212  is carried upstream to premix chamber  211 . For clarity purposes, this region is enlarged and shown in FIG.  9 . Air used for effusion cooling is channeled into premix chamber  211  through multiple rows of angled film cooling holes  217  where it forms a cooling film along liner shell  201  and due to its high velocity, penetrates the boundary layer to mix with the previously discharged fuel and air prior to combustion. Cooling holes  217  are angled such that air entering the combustor from the holes is directed towards the combustion chamber. These angled film cooling holes may also be angled tangentially with respect to the combustor centerline to impart a swirling component to the cooling air as explained below (see FIG.  10 ). Additional air is introduced to the premix chamber through dilution holes  218  for mixing with the upstream fuel and air. The remaining air travels upstream to the forward end of liner  201  and is introduced through one of four regions. Air can enter premix zone  211  through multiple rows of angled film cooling holes  232  in dome plate  216 . These holes may be angled in a tangential direction relative to the combustor centerline to impart swirl in the cooling air. A portion of the air dedicated for dome plate  216  is used to cool the nozzle tubes  234  by way of impingement cooling. Cooling air impinges upon the backside of nozzle tubes  234  through impingement holes  233  and is then directed downstream into premix chamber  211 . The second air route is through nozzle tube  234  located within aperture  213 . The air passing through this region travels through the fuel nozzle air swirler (not shown) where it is premixed with fuel prior to entering premix chamber  211 . The third passage for air entering the combustion system through the dome cap assembly is through an inner substantially cylindrical tube  214  and swirler  215 , which is located within inner tube  214 . The swirler is comprised of inner and outer cylindrical tubes  237  and  238 , respectively. Joining these concentric tubes is an array of angled vanes  239 . This air will mix with the fuel and air of the secondary fuel nozzle (not shown) and exit into the secondary combustion chamber  210 . The fourth method and structure for introducing air into the premix chamber is through cavity  235  formed by inner tube  214  and an outer tube  236 , which are co-axial. Air exits channel  235  through multiple rows of angled film cooling holes  217  in outer center tube  236  or through an aft swirler  231 , which is co-axial to swirler  215 , and discharges the air into secondary combustion chamber  210 . Again, the angled holes  217  direct cooling air towards the combustion chamber and may be angled circumferentially as well, depending upon the application. 
     The premix chamber liner shell  201  cooling hole pattern utilized on the present invention is shown in greater detail in a cross-sectional view of premix chamber liner shell  201  as in FIG.  10 . The cooling holes  217  (all the holes shown in FIG. 10 except hole  218 ) are angled both towards the combustion chamber and circumferentially in order to increase cooling surface area and to induce swirl within the premix chamber, hence improving fuel and air mixing, which will result in lower emissions. Typical combustor liner wall thickness for effusion cooling is thicker than combustors shown in the prior art, with wall thickness at a minimum of 0.0625″, ranging up to 0.25″. The liner wall is laser drilled with a specified pattern of cooling holes  217  and dilution holes  218 . Typical effusion cooling hole diameters EH can range from 0.015″ to 0.125″, while dilution hole DH diameters can range from 0.4″ to 1.5″ as well. Diffusion hole  218  is typically drilled normal to liner shell  201 , while effusion holes  217  are drilled at an angle A relative to the premix chamber centerline axis, where angle A can range from 15 deg to 60 deg from centerline. Drilling these holes at such an angle will result in cooling hole length L, which is a function of angle A. In order to induce swirl within the combustor, which will improve overall mixing of fuel and air, the cooling holes are also drilled at a circumferential angle B, which typically ranges from 15-60 degrees. The cooling holes are spaced apart a circumferential distance C and an axial distance S. These distances are specifically calculated depending upon the application and operating conditions to ensure that the proper amount of cooling air is applied to liner shell  201  for cooling purposes. 
     Operation of the dual-stage, dual-mode combustor disclosed in the present invention is similar to those of similar configuration where ignition is established in the primary zone, or premix chamber, first. Upon confirmation of a steady flame in primary or premix zone  211 , the fuel circuits are opened to the secondary fuel system (not shown) located within center body  214  and flame is established in secondary combustion chamber  210 , aft of venturi throat  219 . Upon confirmation of flame in the secondary combustion chamber  210 , the fuel supply is gradually reduced to the primary fuel nozzles in nozzle tubes  234  until the flame is extinguished, while fuel supply to the secondary fuel system (not shown) is increased in order to transfer all flame to the secondary combustion chamber  210 . Fuel supply is gradually increased to the primary fuel nozzles (not shown) to create a premix of fuel and air in premix chamber  211  while fuel to the secondary system is decreased. This premix fuel and air in the primary premix chamber travels downstream to the secondary combustion chamber where ignition occurs. 
     The benefits to the present invention are numerous over similar hardware configurations. The angled cooling holes in premix chamber liner  201  provide for improving the film cooling effectiveness along the liner skin as well as allowing the cooling air to penetrate the gas path and mix more completely with the fuel within the premix chamber. This configuration is advantageous for an industrial application where increased weight from thicker liner walls is not a primary concern but improved emission is critical. The angled cooling holes by design do not require as much cooling air, so air originally designated for liner cooling can now be introduced further upstream in the premixing process. This extra air introduced further upstream in the liner pushes the fuel/air ratio lower and lowers the flame temperature by allowing for a longer mixing period, hence more complete mixing. These are both key elements that lower NOx levels. A further element to lower NOx emissions of the liner is to introduce the venturi cooling air into the combustion process, which will further reduce the fuel to air ratio and flame temperature, again lowering the resulting NOx levels. This can be accomplished by utilizing the improvements disclosed in U.S. patent application Ser. No. 09/605,765 entitled “Combustion Chamber/Venturi Cooling for a Low NOx Emission Combustor” assigned to the same assignee as the present invention. By reintroducing the cooling air from the venturi into the combustion process in combination with improving the upstream mixing pattern and increasing air flow into the premixing process, overall NOx emissions are substantially reduced. 
     The instant invention has been shown and described herein in what is considered to be the most practical and preferred embodiment. It is recognized, however, that departures may be made therefrom within the scope of the invention and that obvious modifications will occur to a person skilled in the art.