Patent Publication Number: US-2015059349-A1

Title: Combustor chamber cooling

Description:
TECHNICAL FIELD 
     The application relates generally to gas turbine engines and, more particularly to cooling a combustor for such engines. 
     BACKGROUND 
     In aviation gas turbine engines, the amount of air supplied for combustion and dilution may be optimized for operability and to minimize emissions such as nitrous oxide (NOx), carbon monoxide (CO), hydrocarbons (HC), etc. Therefore, it is often desirable that the amount of air supplied for cooling combustor walls be minimized, which poses challenges to meeting the durability requirements of the combustor walls, because the reduction in combustion wall cooling air may lead to unwanted material oxidation, thermal mechanical fatigue and/or thermal wall buckling due to thermal gradients. Particularly in small aero gas turbine engines, the total amount of air available for combustor wall cooling within the gas turbine thermodynamic cycle can be limited, especially where rich-burn combustion is sought. Therefore it is a challenge to optimize the combustor wall cooling while still meeting the durability requirements of the combustor. 
     SUMMARY 
     There is provided a combustor for a gas turbine engine comprising: inner and outer liners spaced apart from each other to define a chamber therebetween including a primary zone adjacent an upstream end of the chamber and a secondary zone downstream of the primary zone, the upstream end being closed by a dome, a dome heat shield disposed within the chamber, adjacent and spaced apart from the dome, an inner liner heat shield disposed within the chamber, adjacent and spaced apart from the inner liner, an outer liner heat shield disposed within the chamber, adjacent and spaced apart from the outer liner, the inner and outer liner heat shields extending from the upstream end of the chamber over the primary zone and terminating at the secondary zone, the inner and outer liners and the inner and outer liner heat shields defining a plurality of apertures therein for introducing dilution air jets into the chamber, the combustor including a plurality of impingement holes defined in the dome and inner and outer liners for directing impingement cooling air therethrough and into the combustor to impinge on a cold side of the respective heat shields, and a discharging apparatus to direct exhausted impingement cooling air discharged from the cold side of the respective heat shields to flow along and substantially parallel to a hot side of the respective inner and outer liner heat shields, the combustor further including a plurality of effusion holes defined in a section of the inner and outer liners free of coverage by the respective inner and outer liner heat shields, the effusion holes directing effusion cooling air therethrough and into the combustor to effusion cool said section of the inner and outer liners. 
     There is also provided a method for hybrid cooling a combustor of a gas turbine engine comprising steps of: a) providing heat shield panels in a chamber of the combustor to protect respective upstream sections of inner and outer liners of the combustor, the upstream sections of the inner and outer liners substantially covering a primary zone of the chamber; b) directing impingement cooling air for impingement cooling a cold side of the heat shield panels and then discharging a first portion of the impingement cooling air along and substantially parallel to a hot side of the heat shield panels and a second portion of the impingement cooling air into areas downstream of the heat shield panels, to thereby have the discharged cooling air substantially outside of the primary zone in order to reduce emission formation; and c) directing cooling air into a downstream area of the chamber for effusion cooling of respective downstream sections of the inner and outer liners free of coverage by the heat shield panels. 
    
    
     
       DESCRIPTION OF THE DRAWINGS 
       Reference is now made to the accompanying figures in which: 
         FIG. 1  is a schematic side cross-sectional view of a gas turbine engine; 
         FIG. 2  is a chart illustrating a relationship between flame temperature and fuel air ratio of gas turbine engines; 
         FIG. 3  is a partial cross-sectional view of the combustor of the gas turbine engine of  FIG. 1 , illustrating cooling arrangements of the combustor according to one embodiment of the present disclosure; and 
         FIG. 4  is an enlarged portion of the combustor of  FIG. 3 , showing an upstream end of the combustor chamber with detailed features. 
     
    
    
     It will be noted that throughout the appended drawings, like features are identified with like reference numerals. 
     DETAILED DESCRIPTION 
       FIG. 1  illustrates an aircraft turbofan gas turbine engine presented as an example of the application of the described subject matter, including a housing or nacelle  10  including an annular core casing or engine outer case  13 , a low pressure spool assembly seen generally at  12  which includes a fan assembly  14 , a low pressure compressor assembly  16  and a low pressure turbine assembly  18 , and a high pressure spool assembly seen generally at  20  which includes a high pressure compressor assembly  22  and a high pressure turbine assembly  24 . The annular outer case  13  surrounds the low and high pressure spool assemblies  12  and  20  in order to define a main fluid path (not numbered) therethrough. A combustor  26  is provided in the main fluid path. The combustor  26  according to this embodiment is a reverse-flow type as illustrated. However, the subject matter described below is also applicable to a combustor of a straight-flow type. 
     In small aero gas turbine engine combustors, the total amount of air available for combustor wall cooling is very limited especially with rich burn technologies. To improve small aero engine Specific Fuel Consumption (SFC), the pressure drop across the combustor wall has to be reduced below 3.0%, whereas in larger engines a pressure drop &gt;3.0% across respective inner and outer liners and &gt;4.5% pressure drop in a dome panel bulkhead due to optimal wall cooling may be possible. The SFC of larger engines can be improved by other means, such as ultra high pressure ratio and active turbine tip clearances etc., which are more affordable due to engine size. 
     In smaller engines, however, to improve emissions more mass flow is often desirable for combustion. Due to various constraints such as compressor mass flow, pressure ratios, combustor delta pressure, and a relatively larger combustor surface area for cooling, the wall cooling can be optimized in order to increase mass air flow for the combustion necessary to improve emissions. 
     To minimize NOx formation and to reduce the flame temperature in a primary zone of the combustor for high or medium pressure ratio engines, a high Fuel Air Ratio (FAR&gt;0.1, Stoichiometric FAR=0.068) is desirable. As illustrated in  FIG. 2 , these pockets of high Fuel Air Ratio (FAR) in a primary zone and advance effusion cooling with coolant, would reduce the FAR toward the stoichiometric (as indicated by circles A and B in  FIG. 2 ) and would stabilized the flame near the wall. This will increase NOx formation due to the high local gas temperatures at close to the stoichiometric FAR. Carbon monoxide (CO) and hydrocarbon (HC) formation can be reduced if the wall temperature in the primary zone is kept relatively high. Circle C in  FIG. 2  indicates a Fuel Air Ratio (FAR) operating zone of low pressure ratio engine combustors. Additional effusion coolant would decrease the local FAR and move same away from the stoichiometric region, as indicated by circle D in  FIG. 2 , thereby further reducing the flame temperature. 
     Therefore, to optimize the combustor cooling to improve emission reduction and durability of the combustor, the combustor may be provided with hybrid cooling techniques using heat shields in an upstream end of the combustor chamber with impingement cooling and having exhausted impingement cooling air substantially outside the primary zone in the chamber, meanwhile using effusion cooling for the single-skin wall of a downstream section of the combustor chamber which includes therein the secondary zone. 
     According to one embodiment of the present disclosure as shown in  FIG. 3-4 , the combustor  26  which may be used with rich burn techniques for a gas turbine engine of high/medium pressure ratio, includes annular inner and outer liners  30 ,  32  spaced apart from each other to define an annular chamber  28  therebetween. The annular chamber  28  may have a straight upstream section with an upstream end closed by a bulkhead or dome  34 . The dome may be formed with a plurality of dome panels. The annular chamber  28  may include a downstream section in a curved shape (as in a reverse type of combustor) with an open downstream end for discharging combustion gases to power the turbines  24 ,  18  of the engine. A plurality of circumferentially spaced apart fuel nozzles  36  with swirlers  38  are attached to the dome  34  for introduction of fuel and air into the chamber  28  for combustion. Combustion in the chamber  28  may principally take place in the primary zone  40 , adjacent the upstream end of the chamber  28 , and the combustion reaction may then continue in the secondary zone  42  downstream of the primary zone  40 . 
     A dome heat shield  44 , which may be formed with a plurality of dome heat shield panels, is disposed within the chamber  28 , adjacent and spaced apart from the dome  34 , attached to the dome  34  to protect the dome from exposure to the hot combustion gases in the chamber  28 . An inner liner heat shield  46  which may be formed with a plurality of inner liner heat shield panels, is disposed within the chamber  28 , adjacent and spaced apart from the inner liner  30 , attached to the inner liner  30  to protect the inner liner  30  from being exposed to the hot combustion gases within the chamber  28 . An outer liner heat shield  48  which may be formed with a plurality of outer liner heat shield panels, is disposed within the chamber  28 , adjacent and spaced apart from the outer liner  32 , and is attached to the outer liner  32  to protect the outer liner from being exposed to the hot combustion gases within the chamber  28 . The inner and outer liner heat shields  46  and  48  extend from the upstream end of the chamber  28  over the primary zone  40  and terminate at the secondary zone  42  and therefore are used only to protect the upstream sections of the inner and outer liners  30  and  32 . Each of the dome heat shield  44  and inner and outer liner heat shields  46 ,  48  may include a plurality of pin fins  49  projecting from a cold side thereof (facing the respective dome  34 , inner and outer liners  30 ,  32 ) to increase contact areas with cooling air to increase heat transfer in convection cooling. 
     The inner and outer liners  30 ,  32  and the inner and outer liner heat shields  46 ,  48  define a plurality of apertures therein which are in communication with the chamber  28 , for introducing dilution air jets  52  into the chamber  28  between the primary zone  40  and the secondary zone  42 . The dilution air helps to reduce flame temperature by quenching in the secondary zone  42  and provides for combustor exit temperature distribution acceptable for turbines  24 ,  18 . 
     A plurality of impingement cooling holes  54  (see  FIG. 4 ) may be provided in the respective inner and outer liners  30 ,  32 , for example at a relatively upstream section, as well as in the dome  34  for directing impingement cooling air therethrough and into the combustor  26  to impinge on the cold side of the respective dome heat shield  44  and liner heat shields  46  and  48 . The impingement cooling air flows in the space between the dome  34  and the dome heat shield  44 , between the inner liner  30  and the inner liner heat shield  46  and between the outer liner  32  and the outer liner heat shield  48 , contacting the cold side of the respective heat shields  44 ,  46 ,  48  and the pin fins  49 , therefore convection cooling also takes place. 
     A discharging apparatus may be provided for directing exhausted impingement cooling air discharged from the cold side of the respective heat shields to flow substantially along and parallel to a hot side (facing the hot combustion gases in the chamber) of the respective heat shields. For example, the dome shield  44  is positioned with respect to the inner and outer liner heat shields  46 ,  48  to provide respective gaps  56 ,  58  between the dome heat shield  44  and the inner liner heat shield  46 , and between the dome heat shield  44  and the outer liner heat shield  48 . The dome heat shield  44  is configured such that the impingement cooling air introduced to the space between the dome heat shield  44  and the dome  34  is forced to flow towards the respective gaps  56 ,  58  and is discharged therefrom to form a cooling film along and substantially parallel to the hot side to cool the hot side of the respective inner and outer liner heat shields  46 ,  48 . The exhausted impingement cooling air discharged from the respective gaps  56 ,  58  in the form of a cooling air film along the hot side of the respective inner and outer liner heat shields  46 ,  48  is substantially outside the primary zone  40  to reduce the NOx, CO and HC emission formation in the combustion gases. Thus, there is no low momentum effusion to be discharged from the exhausted impingement cooling air into the pocket of high fuel air ratio as indicated in circle A of  FIG. 2 , which helps to reduce the emissions and to maintain low gas temperature near the inner and outer liner heat shields  46  and  48  as discussed above with reference to  FIG. 2 . 
     A portion of exhausted impingement cooling air in the respective spaces between the inner liner  30  and the inner liner heat shield  46 , and between the outer liner  32  and the outer liner heat shield  48  flows towards the upstream end of the chamber  28  (as indicated by the arrows shown in  FIG. 3 ) to be discharged from the respective gaps  56 ,  58  to join the cooling air film along and substantially parallel to the hot side of the respective inner and outer liner heat shields  46 ,  48 . Another portion of exhausted impingement cooling air in the respective spaces between the inner liner  30  and the inner liner heat shield  46 , and between the outer liner  32  and the outer liner heat shield  48  flows towards the downstream section of the chamber  28  and discharged from such spaces at respective downstream ends of the inner and outer liner heat shields  46 ,  48 . 
     Optionally, the discharging apparatus may include a plurality of splash louvers on the cold side at the downstream end of the respective inner and outer liner heat shields  46 ,  48 , acting as a film starter for the portion of the exhausted impingement cooling air discharged from the downstream end of the inner and outer liner heat shields  46 ,  48  to flow along a downstream section of the respective inner and outer liners  30 ,  32 . 
     Optionally, the inner and outer liner heat shields  46 ,  48  may be provided with two rows of small effusion hose  60 , located downstream of the apertures  50  for introduction of dilution jet air, near the downstream end of the inner and outer liner heat shields  46 ,  48 . Therefore, a further portion of the exhausted impingement cooling air is discharged from the effusion hose  60  and may enter the secondary zone  42  (see  FIG. 3 ). 
     The swirlers  38  are provided with swirl air passages  62  for introduction of air flows in a swirling flow into the chamber  28  to mix with the fuel ejected by the fuel nozzles  36  for combustion. In this embodiment, the swirlers  38  each may include cooling air passages  64  for introducing cooling air into the combustor independent from the air flow introduced by the swirl air passages  62  of the swirlers  38 . Each swirler  38  is configured as to provide a cooling air director  66  for directing the cooling air introduced from the cooling air passages  64  to generate a cooling air film along and substantially parallel to a hot side of the dome heat shield  44  to cool the same. 
     In  FIG. 3 , the downstream section of the inner and outer liners  30 ,  32  are made of single skin without the coverage of any heat shields. A plurality of effusion holes  68  are defined in this downstream section of the inner and outer liners  30 ,  32  for directing cooling air therethrough and into the chamber  28  to effusion cool the downstream section of the inner and outer liners  30 ,  32 , which is an efficient cooling method. 
     It is noted that in the combustor  26  of the reverse flow type, the downstream section of the inner and outer liners  30 ,  32  are made of curved wall of large exit duct (LED) and small exit duct (SED) to turn the flow. It is very cumbersome to manufacture curved heat shields and difficult to minimize leakage between heat shields and the outer impingement skin of the combustor inner and outer liners. The downstream section of the inner and outer liners substantially define the secondary zone therebetween in which most of the fuel is already oxidized and there is very low fuel air ratio, and therefore the effusion flow entering the effusion holes  68  and into the secondary zone will not help to form NOx, CO and HC emission. 
     The hybrid cooling techniques for a gas turbine engine combustor, such as using rich burn techniques, minimize cooling air entering in the primary zone and thus high FAR can be maintained therein. Lack of effusion air in the primary zone with high FAR flame will not be stabilized locally near the combustor liners. This also helps to minimize thermal gradients in the structure members of the combustor liners by eliminating direct exposure to the flame in the primary zone, thus minimizing combustor liners buckling. The hybrid cooling techniques improve the reduction of NOx at high engine power conditions and CO/HC emission at low engine power conditions. 
     The above description is meant to be exemplary only, and one skilled in the art will recognize that changes may be made to the embodiments described without departing from the scope of the described subject matter. Modifications which fall within the scope of the described subject matter will be apparent to those skilled in the art, in light of a review of this disclosure, and such modifications are intended to fall within the appended claims.