Abstract:
There is provided a method for improving the combustion efficiency of a combustor of a gas turbine engine powering an aircraft. The method comprises selectively using two distinct fuel injection units or a combination thereof for spraying fuel in a combustion chamber of the combustor of the gas turbine engine. A first one of the two distinct fuel injection units is selected and optimized for high power demands, whereas a second one of the two distinct fuel injection units is selected and optimized for low power level demands. In operation, the fuel flow ratio between the two distinct injection units is controlled as a function of the power level demand.

Description:
RELATED APPLICATIONS 
       [0001]    This application is continuation of U.S. patent application Ser. No. 13/915,990, filed Jun/. 12, 2013, which is a divisional of U.S. patent application Ser. No. 13/071,997 filed on Mar. 25, 2011, now U.S. Pat. No. 8479492, issued Jul. 9, 2013, the content of which is hereby incorporated by reference. 
     
    
     TECHNICAL FIELD 
       [0002]    The application relates generally to gas turbine engines and, more particularly, to a hybrid system for injecting fuel into a combustor. 
       BACKGROUND OF THE ART 
       [0003]    Gas turbine engines used for powering aircrafts comprise a combustor in which fuel is mixed with compressed air and ignited to provide combustion gases for the turbine section of the engine. In a slinger combustion system, fuel is delivered and atomized through spraying fuel through a rotary fuel slinger. The rotary fuel slinger is designed for maximum fuel flow and optimized for cruise condition to improve the combustion efficiency and thus reduce smoke and gaseous emission. Thus at low power levels, when the slinger rotates at lower speeds, fuel tends to not atomize properly, thereby resulting in low combustion efficiency, and high emission/smoke/particulates/ unburned hydrocarbons. 
         [0004]    Conventional rotary slingers have to be operated at high speed for properly atomizing the fuel. When, the slinger is rotated at low speeds, such as during starting and altitude relight conditions, the fuel atomization effect of the slinger is relatively poor, thereby requiring a relatively expensive and complex architecture for the ignition system with relatively long igniters to deliver spark energy close to the stinger system. Starting a slinger combustor at low speeds and at high altitudes without relatively complex high pressure fuel injection system has heretofore been challenging. 
       SUMMARY 
       [0005]    In one aspect, there is provided a hybrid slinger combustor system for an aero gas turbine engine powering an aircraft, the combustor system comprising a combustor shell defining a combustion chamber, the combustion chamber having first and second combustion zones; two distinct fuel injector units for respectively spraying fuel into said first and second combustion zones, said two distinct fuel injector units including a rotary fuel slinger for spraying fuel radially outwardly into the first combustion zone, and a set of circumferentially spaced-apart fuel nozzles for spraying fuel into the second combustion zone; and a control unit controlling the rate of fuel flow to said rotary fuel slinger and said set of fuel nozzles as a function of the power demand of the gas turbine engine. 
         [0006]    In a second aspect, there is provided a method for improving the combustion efficiency of a combustor of a gas turbine engine powering an aircraft, comprising: selectively using two distinct fuel injection units or a combination thereof for spraying fuel in a combustion chamber of the combustor of the gas turbine engine, a first one of the two distinct fuel injection units being selected and optimized for high power demands, whereas a second one of the two distinct fuel injection units being selected and optimized for low power level demands, and controlling a fuel flow ratio between said two distinct injection units as a function of the power level demand. 
     
    
     
       DESCRIPTION OF THE DRAWINGS 
         [0007]    Reference is now made to the accompanying figures, in which: 
           [0008]      FIG. 1  is a schematic cross-sectional view of a turbofan gas turbine engine; 
           [0009]      FIG. 2  is a schematic cross-sectional view of the combustor section of the gas turbine engine, the combustor section having a hybrid slinger combustion system including a high power combustion zone supplied with fuel by a slinger and a low power combustion zone supplied with fuel by a set of fuel nozzles; and 
           [0010]      FIGS. 3   a  to  3   c  are graphic representations illustrating the fuel flow distribution between the slinger and the fuel nozzles at different power level conditions. 
       
    
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS 
       [0011]      FIG. 1  illustrates a turbofan gas turbine engine  10  of a type preferably provided for use in subsonic flight, generally comprising in serial flow communication a fan  12  through which ambient air is propelled, a multistage compressor  14  for pressurizing the air, a combustor  16  in which the compressed air is mixed with fuel and ignited for generating an annular stream of hot combustion gases, and a turbine section  18  for extracting energy from the combustion gases. 
         [0012]    As can be appreciated from  FIG. 2 , the combustor  16  is a hybrid slinger combustor combining two distinct fuel injector units, a rotary fuel slinger  20  and a set of spaced-apart fuel nozzles  220  As will be discussed in further details hereinbelow, the rotary fuel slinger  20  may be optimized for high power engine demands, such as a during take-off and climb phases of a flight, while the set of individual fuel nozzles  22  may be optimized for low power engine demands, for example, at ground or flight idle. Under certain flight conditions, such as at cruise power level, the two distinct fuel injector units, the rotary fuel slinger  20  and the set of fuel nozzles  22 , may be both used to co-inject fuel according to a predetermined fuel flow ratio. 
         [0013]    Referring more particularly to  FIG. 2 , it can be appreciated that the combustor  16  is housed in a plenum  25  supplied with compressed air from the compressor diffuser  27  of the compressor  14 . The combustor  16  has an annular combustor shell  24  concentrically mounted about the engine centerline  11  in the plenum  25 . The combustor shell  24  may have a front annular liner  26  and a rear annular liner  28 . The front and rear annular liners  26  and  28  are axially spaced-apart to define therebetween a combustion chamber  30 . As schematically depicted by flow arrows  32 , the front and rear liners  26  and  28  each include a plurality of air inlet openings for allowing air to flow from the plenum  25  into the combustion chamber  30 . Cooling holes (not shown) such as effusion cooling holes, may also be defined in the front and rear liners  26  and  28  to provide cooling to the liners  26  and  28 . 
         [0014]    As schematically shown in  FIG. 2 , the rotary fuel slinger  20  is mounted for rotation with an engine shaft  34  coupled to the compressor or the turbine rotor. The rotary fuel slinger  20  is axially aligned with a radially inner circumferential opening  36  defined in the combustor shell  24 . The rotary fuel slinger  20  is configured to atomize and spray fuel radially outwardly through the circumferential opening  36  into a first combustion zone  38  of the combustor chamber  30 . A fuel manifold  40  extends into the plenum  25  for directing a flow of fuel from a fuel source (not shown) to the rotary fuel slinger  20 . As the slinger rotates  20 , fuel is centrifuged through outlet holes  42  defined in the slinger  20 , thereby atomizing the fuel into tiny droplets and evenly distributing the fuel into the first combustion zone  38  of the combustor chamber  30 . 
         [0015]    The set of individual fuel nozzles  22 , which may be of any suitable types, are uniformly circumferentially distributed about the combustions chamber  30  and disposed generally downstream of the rotary slinger  20  relative to the flow of combustion gases through the combustions chamber  30 . By way of example, the set of fuel nozzles  22  may be composed of three or four air assisted fuel nozzles (low pressure fuel system). The fuel nozzles  22  extend in respective openings defined in the front liner  26  of the combustor shell  24  and are disposed to spray fuel into a second combustion zone  44  of the combustion chamber  30 . The fuel nozzles are connected to the fuel source via any appropriate fuel manifold structures (not shown). The fuel nozzles manifold can be integrated to the slinger fuel manifold. Valves (not shown) may be provided to control the split of fuel flow between the slinger  20  and the fuel nozzles  22 . 
         [0016]    An appropriate number of igniters (only one being schematically shown in  FIG. 2  at  46 ) are provided to ignite the fuel supplied by both the slinger  20  and the fuel nozzles  22 . The igniters  46  may all be disposed to provide spark energy in the second combustion zone  44  only. By using the fuel nozzles  22  in place of the fuel slinger  20  at ground or flight idle, it is possible to eliminate the need for long igniters which are typically required to deliver spark energy very close to conventional slinger systems in order to compensate for the poor atomization provided by the slinger when operated at low rotational speeds. Due to lower number of fuel nozzles, the fuel nozzles internal cavities can be designed to minimize internal carbon formation in addition to the optimized fuel atomization. The fuel nozzle tip orifice and internal passages may be higher than certain size to minimize internal carbon formation on the wall. 
         [0017]    As mentioned above, the rotary fuel slinger  20  is suited for high power conditions (e.g. take-off, climb and cruise power levels). The fuel nozzles  22  are mainly used for improved starting/altitude relight and other low power level conditions. The fuel nozzles  22  provide for better fuel atomization than the fuel stinger  20  when the engine  10  is operated at low power levels. Such a hybrid or dual mode injection system allows optimizing a first one of the dual fuel injectors for low power fuel consumption and a second one of the injectors for high power fuel consumption. This provides for improved combustion efficiency and lower smoke emission as compared to conventional slinger combustors. 
         [0018]    The split of fuel flow between the rotary fuel slinger  20  and the fuel nozzles  22  is controlled by a control unit  50 . The control unit  50  is configured for controlling the flow of fuel to the rotary fuel slinger  20  and the fuel nozzles  22  as a function of the power demand. 
         [0019]      FIGS. 3   a  to  3   c  graphically illustrate three possible fuel schedules for the hybrid slinger combustions system, each graph illustrating the relative use of the stinger  20  and the set of fuel nozzles  22  in terms of fuel flow during ground operation and various phases of flight, including: ground idle, take-off, climb, cruise and decent. 
         [0020]    According to the first option illustrated in  FIG. 3   a , at ground idle, the fuel is solely injected into the combustion chamber  30  by the fuel nozzles  22 . The fuel flow through the fuel nozzles  22  at ground idle is about 20% to about 35% of the maximum fuel flow (Le. the take-off fuel flow). The slinger  20  only starts injecting fuel into the combustion chamber  30  during the ground idle to take-off acceleration phase. At the same time, the nozzle fuel flow is reduced to zero. The flow of fuel through the fuel nozzles  22  remains at zero during the various flight phases, including the climb and cruise phases. During flight all the fuel is atomized through the rotary fuel slinger  20 . The fuel slinger  20  is thus the primary fuel injector during the flight. At the decent approach, the fuel flow is switched hack to the fuel nozzles  22  as during the first ground idle phase of the engine operation. 
         [0021]      FIG. 3   b  illustrates a second option in which the fuel nozzles  22  atomise a small portion (e.g. 10%) of the fuel required during flight. According to this scenario, during flight the fuel nozzles  22  will have fuel just enough to maintain a flame. The amount of fuel through the rotary fuel slinger  20  during flight will total the required amount of fuel minus the fuel flowing through the fuel nozzles  22 . 
         [0022]      FIG. 3   c  illustrates a third option in which through out the engine running, the fuel nozzles  22  will have the ground idle fuel flow condition (i.e. the fuel flow will remain constant at about 30% to 35% of the maximum fuel flow). Again, the fuel will be supplied to the rotary slinger  20  at the beginning of the ground idle to take-off acceleration phase. During flight, the slinger fuel flow will total the required fuel flow minus the fuel through the fuel nozzles  22  (the ground idle fuel flow). 
         [0023]    As can be appreciated from the description of  FIGS. 3   a  to  3   c , the fuel flow ratio between the rotary slinger  20  and the fuel nozzles  22  is controlled by the control unit  50  as a function of the variation of the power demand over a full range of engine power settings. 
         [0024]    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 invention disclosed. Modifications which fall within the scope of the present invention 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.