Abstract:
A system and method for modifying the supply of fuel to injectors to attenuate combustion oscillations in a gas turbine engine. The gas turbine engine may comprise a combustor, a plurality of injectors and a manifold. The plurality of injectors may be operable to provide fuel to the combustor. The manifold may be configured to supply fuel to all of the plurality of injectors or to only a portion of the plurality of injectors in reaction to a determination of an existence of combustion oscillations.

Description:
RELATION TO OTHER PATENT 
     This application claims the benefit of prior provisional patent application Ser. No. 60/710,116, filed Aug. 22, 2005. 
    
    
     TECHNICAL FIELD 
     The present disclosure relates generally to attenuating combustion oscillations in gas turbine engines, and more specifically to a system and method of modifying the supply of fuel to injectors to attenuate combustion oscillations. 
     BACKGROUND 
     Producers of gas turbine engines have made great strides in reducing regulated emissions such as NOx. However, these strides have led to various instabilities in combustion, such as combustor thermo-acoustic oscillations. This problem may be brought about by the coupling of the heat release and pressure waves, which produce a resonance with a characteristic frequency usually corresponding to one or more natural frequencies of the combustion chamber. This has been historically described by the well-known Rayleigh Mechanism. Such oscillations in the combustor may result in mechanical and thermal fatigue to combustor hardware which may lead to other operational problems that may have adverse affects on the engine. 
     Several attempts have been made to eliminate, diminish or prevent thermo-acoustic oscillations. One such attempt to attenuating oscillations was to decouple the heat release form the pressure wave by moving the fuel introduction point in the injector so that the residence time to the flame was different than that required to sustain resonant oscillations. This may be accomplished by moving the fuel spokes along the length of the injector main fuel and air flow path. 
     Another attempt to attenuate oscillations was to introduce airflow in a row of holes around the circumference of the injector barrel. The axial location of the row of holes along the barrel was determined so as to dilute the fuel to air ratio in an attempt to provide a non-sinusoidal variation in the energy input to the flame. However, problems with such attempts has been in the limited number of frequencies of oscillations that could be attenuated. Unfortunately, some engines exhibit more than one frequency of oscillation when they are at different power settings. 
     The system and method of the present disclosure is set forth to overcome at least one of the problems described above. 
     SUMMARY OF THE INVENTION 
     It is to be understood that both the foregoing general description and the following detailed description are exemplary and explanatory only and are not restrictive of the invention as claimed. 
     In one aspect of the present disclosure, a gas turbine engine is disclosed. The gas turbine engine comprises a combustor, a plurality of injectors and a manifold. The plurality of injectors may be operable to provide fuel to the combustor. The manifold may be configured to supply fuel to all of the plurality of injectors or to only a portion of the plurality of injectors in reaction to a determination of an existence of combustion oscillations. 
     In another aspect of the present disclosure, a method of controlling combustion oscillations on a gas turbine engine is disclosed. The method may comprise the steps of determining an existence of a combustion oscillation and preventing fuel to at least one of a plurality of injectors. 
     In another aspect of the present disclosure, a system for controlling fuel delivery is disclosed. The system comprises a sensor, a manifold and at least one valve. The sensor may be configured to determine an existence of combustion oscillations. The manifold may be configured to supply fuel to all of a plurality of injectors or to only a portion of the plurality of injectors. The at least one valve may be operable to control fuel to at least a portion of the manifold in reaction to the determination of the existence of combustion oscillations. 
     In another aspect of the present disclosure, a method of attenuating combustion oscillations is disclosed. The method may comprise the steps of determining an existence of a combustion oscillation within a combustion system and controlling the combustion oscillations by creating a non-homogenous temperature condition within a combustion zone of the combustion system. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the invention and, together with the description, serve to explain the principles of the invention. In the drawings, 
         FIG. 1  illustrates a cross-sectional view of a gas turbine engine with a combustor and injector; 
         FIG. 2  illustrates a typical combustor/injector configuration; 
         FIG. 3  illustrates a control circuit for controlling pilot fuel; 
         FIG. 4  illustrates a graph showing machine vibration of a gas turbine engine in normal operation under full load; 
         FIG. 5  illustrates a graph showing machine vibration of a gas turbine engine in operation with the present invention under full load; 
         FIG. 6  illustrates a graph showing machine vibration of a gas turbine engine in without the present invention under part load; and 
         FIG. 7  illustrates a graph showing machine vibration of a gas turbine engine in operation with the present invention under part load. 
     
    
    
     DETAILED DESCRIPTION 
     Reference will now be made in detail to embodiments of the invention, examples of which are illustrated in the accompanying drawings. Whenever possible, the same reference numbers will be used throughout the drawings to refer to the same or like parts. 
     Referring to  FIG. 1 , a gas turbine engine  10  is shown, but not in its entirety. The gas turbine engine  10  includes an airflow delivery system  12  for providing combustion air and for providing cooling air for cooling components for the engine  10 . The engine  10  includes a turbine section  14 , a combustor section  16  and the compressor section  18  operatively connected to the turbine section  14 . In this application the combustor section  16  includes an annular combustion chamber  24  positioned about a central axis  26  of the gas turbine engine  10 . As an alternative, the engine  10  could include a plurality of can combustors without changing the essence of the invention. The annular combustion chamber  24  is operatively positioned between the compressor section  18  and the turbine section  14 . A plurality of fuel injectors  30  (one shown) are positioned in an inlet end portion  32  of the annular combustion chamber  24 . The turbine section  14  includes a first stage turbine  34  being centered about the central axis  26 . 
     An annular combustion zone  38  is enclosed by an inner combustor liner  40  and an outer combustor liner  42  spaced apart a predetermined distance. The inner combustor liner  40  has an inner inlet conical portion  44  and an inner outlet conical portion  46  axially spaced apart by an inner cylindrical liner portion  48 . The inner inlet conical portion  44  connects with fuel injector  30  in a normal fashion. The inner outlet conical portion  46  terminates proximate the turbine section  14 . While the combustor liners  40  and  42  are shown having multiple pieces, the combustor liners  40  and  42  may also be made from a single piece of conventional high-temperature material without changing the essence of the invention. 
     Similarly the outer combustor liner  42  has an outer inlet conical portion  50  and an outer outlet conical portion  52  axially spaced apart by an outer cylindrical liner portion  54 . The outer inlet conical portion  50  connects in a normal fashion with the fuel injector  30 . The outer outlet conical portion  52  terminates proximate the turbine section  14 . The outer outlet conical portion  52  and the inner outlet conical portion  46  define a combustor outlet nozzle  62 . The combustor outlet nozzle  62  fluidly connects with the turbine section  14 . 
     An outer cooling shield  64  surrounds the outer cylindrical liner portion  54 . The outer cooling shield  64  has a first outer shield portion  66  separated axially from a second outer shield portion  68 . Similarly, an inner cooling shield  80  has a first inner shield portion  82  axially separated from a second inner shield portion  84 . The outer cylindrical liner portion  54  and the outer cooling shield  64 , and the second inner shield portion  84  and the inner cooling shield  80  define cooling air plenums  96  for the annular combustion chamber  24 . A combustion air plenum  94  encircles the annular combustor. 
     The cooling air plenums  96  fluidly connect with the combustion air plenum  94  through an inner cooling air passage  98  and an outer cooling air passage  100 . An outer air passage  102  and inner air passage  104  fluidly connect with the air flow delivery system  12 . The outer air passage  102  and the outer cooling air passage  100  fluidly connect through a plurality of impingement holes (not shown) in the outer cooling shield  64 . Likewise, the plurality of impingement holes fluidly connects the inner cooling air passage  98  with the inner air passage  104 . 
     Referring now to  FIG. 2 , a fuel injector  30  is shown. The fuel injector  30  includes a plurality of passages and a swirler  106 . The passages direct air and fuel into the annular combustion chamber  24 . Pilot fuel and pilot air mix in a pilot passage  108  and exit through a front side  110  of the fuel injector  30 . Additional fuel is added in a premix duct  112  of the fuel injector  30  through orifices  114  in axial alignment with the swirler  106 . The premix duct  112  is an annular passage between an outer barrel  116  of the fuel injector  30  and a center body  118 , downstream of the swirler  106 , in which fuel and air are given ample space and time to mix. Air entering into the fuel injector  30  passes through the swirler  106  and collects fuel provided through the orifices  114 . The swirler  106  features ten axial vanes spanning from the center body  118  to the outer barrel  116 , designed to impose a forced vortex to the air and fuel flow within the premix duct  112 . The air and fuel mix and enter the annular combustor chamber  24  where they ignite. 
     Initially, a pilot flame (not shown) in the annular combustion chamber  24  may be stable, or in other words, the combustion may not be producing undesirable oscillations. Combustion driven oscillations are likely when the changes in heat release are in phase or partial phase with the acoustic pressure disturbances in the combustion chamber  24 . 
     Referring now to  FIG. 3 , a control circuit for controlling pilot fuel is shown. The illustrated control circuit has an air supply  122  and a gas fuel supply  128 . The air supply  122  may feed intermittent pilot shutoff valve solenoid  126  that may actuate an intermittent pilot shutoff valve  142  and primary and secondary gas fuel shutoff valve solenoids  130 , 134  that may actuate primary and secondary gas fuel shutoff valves  132 , 136 . The shutoff valves  126 , 132 , 136  in conjunction with the pilot fuel control valve  146  and the main fuel control valve  150  may control fuel delivery to the manifold  140 . 
     The manifold  140  may be comprised of three independent manifold sections: an intermittent pilot manifold  144  and a continuous pilot manifold  148  for supplying pilot fuel and a continuous main manifold  152  for supplying main fuel. Each manifold section  144 , 148 , 152  may incorporate its own independent fuel inlet feeding the fuel injectors (illustrated in this embodiment by orifices labeled  1 - 12  within each manifold section  144 , 148 , 152 ). In the illustrated embodiment, the shutoff valves  126 , 130 , 134  and control valves  146 , 150  may be operable to allow fuel to the fuel injectors or to deny fuel to the fuel injectors. The continuous pilot manifold  148  may supply pilot fuel to a first portion of the plurality of injectors (labeled  1 , 2 , 5 , 6 , 9 , 10 ), which may always be on. The intermittent pilot manifold  144  may supply pilot fuel to a second portion of the plurality of injectors (labeled  3 , 4 , 7 , 8 , 11 , 12 ), which may also be denied pilot fuel. 
     Although not shown, it is anticipated that the manifold  140  may comprise a variety of sections, each section being controllable to provide fuel to any one, or any number of fuel injectors. 
       FIGS. 4 and 5  illustrate graphs showing machine vibration of a gas turbine engine and show the oscillation amplitude for frequencies between 0 to 1000 Hz under full loads.  FIG. 4  shows the oscillation amplitude without the benefit of the present disclosure.  FIG. 5  shows the oscillation amplitude after implementing the present disclosure. 
       FIGS. 6 and 7  illustrate graphs showing machine vibration of a gas turbine engine and show the oscillation amplitude for frequencies between 0 to 1000 Hz under part loads.  FIG. 6  shows the oscillation amplitude without eh benefit of the present disclosure.  FIG. 7  shows the oscillation amplitude after implementing the present disclosure. 
     The oscillation mode was identified using the wave equation in an annular space with non-damping boundary conditions and choked conditions at the turbine inlet defining the downstream end of the control volume. It was determined that creating a non-homogeneous temperature condition around parts of the circumference of the combustion primary combustion zone could be used to ensure that the acoustic velocity in the medium of the primary combustion zone was different than in other circumferential portions of the primary zone. Oscillations are transmitted at acoustic velocity, which is proportional to the square root of the temperature of the combustion gases in the primary combustion zone. Accordingly, creating circumferential zones of differing acoustic velocities would impose a time lag on the oscillation propagation, thus disrupting the period corresponding to the offending oscillation frequency. 
     INDUSTRIAL APPLICABILITY 
     In operation, fuel enters the main fuel line. Air actuated valves  142 , 132 , 136  allow the fuel to flow to the continuous main manifold  152 , the continuous pilot manifold  148 , and the intermittent pilot manifold  144 . The amount of fuel may be controlled using variable or fixed orifices. Under normal operating conditions, fuel will continuously flow to the continuous main manifold  152 , which supplies fuel into the premix duct  112  and into the annular combustion chamber  24 . 
     When oscillations begin, a transducer (not shown) detects excessive amplitudes at the oscillating frequencies and communicates a representative signal to a control module. The control module may shut off fuel to the intermittent pilot manifold  144  and redirect the pilot fuel to the main fuel line. As a result, pilot fuel may only be fed to one-half of the total injectors, but the total fuel may remain substantially constant. The intermittent pilot manifold  144  may remain closed until the oscillations stop, or until the amplitude of the oscillation returns to an acceptable level. 
     The closing of pilot fuel to a portion of the fuel injectors  30  during operation creates a non-homogeneous temperature condition around parts of the circumference of the combustion primary combustion zone. The non-homogeneous conditions may result in different acoustic velocities in regions around the circumference between injectors with fueled pilot lines and injectors with non-fueled pilot lines. Accordingly, creating circumferential zones of differing acoustic velocities impose a time lag on the oscillation propagation, which attenuates the oscillations. 
     The pilot fuel of any of the fuel injectors  30  may be controlled through the manifold. It is envisioned that any number of sequences may be used to interrupt the oscillating frequency. For example, it is envisioned that the pilot line to only one fuel injector  30  may be closed, or a plurality of pilot lines may be closed. However, attenuation of all offending oscillation frequencies may be obtained by shutting off the fuel to the pilots of every other pair of fuel injectors  30  around the annular combustion chamber  24 . For example, in the illustrated embodiment of  FIG. 3 , shutting off injector  3  and  4 ,  7  and  8 , and  11  and  12  proved to eliminate oscillations. It is also envisioned that the oscillating frequency may be interrupted by increasing the fuel to the pilots of a portion of the plurality of injectors. 
     It is noted that the strategy of turning off the pilot fuel to the fuel injectors  30  may also be incorporated full time. In other words, the engine  10  may use all injectors with the pilot fuel lines active during start up to accelerate and bring the engine to a stable operational temperature, thus avoiding temperature distribution into the turbine hot section during the start up sequence, which would occur if only some pilot lines were fueled, especially since a large portion of the total fuel is apportioned to the pilot lines during start up. Post the start up sequence, the pilot lines of alternate pairs of injectors may be permanently turned off in response to any detected combustor oscillations. Fuel from the non-active injector pilot lines may be diverted to the main fuel supply. 
     It will be apparent to those skilled in the art that various modifications and variations can be made in the system and method of the present invention without departing from the scope or spirit of the invention. Other embodiments of the invention will be apparent to those skilled in the art from consideration of the specification and practice of the invention disclosed herein. It is intended that the specification and examples be considered as exemplary only, with a true scope and spirit of the invention being indicated by the following claims and their equivalents.