Abstract:
An Active Combustion Control System and method provides for monitoring combustor pressure and modulating fuel to a gas turbine combustor to prevent combustion dynamics and/or flame extinguishments. The system includes an actuator, wherein the actuator periodically injects pulsed fuel into the combustor. The apparatus also includes a sensor connected to the combustion chamber down stream from an inlet, where the sensor generates a signal detecting the pressure oscillations in the combustor. The apparatus controls the actuator in response to the sensor. The apparatus prompts the actuator to periodically inject pulsed fuel into the combustor at a predetermined sympathetic frequency and magnitude, thereby controlling the amplitude of the pressure oscillations in the combustor by modulating the natural oscillations.

Description:
STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH &amp; DEVELOPMENT 
     This invention was developed with U.S. Government support under U.S. Government Contract No. NAS 3-01135 awarded by the National Aeronautics and Space Administration (NASA) and is subject to the provisions of Section 305 of the National Aeronautics and Space Act of 1958 (42 U.S. C. 2457). The U.S. Government may have certain rights in this invention. 
    
    
     BACKGROUND OF THE INVENTION 
     1. Field of the Invention 
     The present invention relates to controlling combustion in a combustion chamber or combustor. 
     2. Description of the Related Art 
     Air pollution concerns worldwide have led to stricter emissions standards both domestically and internationally. Aircraft emissions are governed by both Environmental Protection Agency (EPA) and International Civil Aviation Organization (ICAO) standards. These standards regulate the emission of oxides of nitrogen (NOx), unburned hydrocarbons (UHC), and carbon monoxide (CO) from aircraft in the vicinity of airports, where they contribute to urban photochemical smog problems. Many aircraft engines are able to meet current emission standards using combustor technologies and theories proven over the past 50 years of engine development. However, stricter engine emissions standards will not be within the capability of current combustor technologies. 
     In general, engine emissions fall into two classes: those emissions formed because of high flame temperatures (NOx), and those emissions formed because of low flame temperatures which do not allow the fuel-air reaction to proceed to completion (UHC &amp; CO). 
     A small window exists where both pollutants are minimized. For this window to be effective, however, the reactants must be well mixed, so that burning occurs evenly across the mixture without hot spots, where NOx is produced, or cold spots, when CO and UHC are produced. Hot spots are produced where the mixture of fuel and air is near a specific ratio when all fuel and air react (i.e. no unburned fuel or air is present in the products). This mixture is called stoichiometric. Cold spots can occur if either excess air is present (called lean combustion), or if excess fuel is present (called rich combustion). 
     Modern gas turbine combustors consist of between 1 and 30 or more mixers, which mix high velocity air with a fine fuel spray. These mixers usually consist of a single fuel injector located at a center of a swirler for swirling the incoming air to enhance flame stabilization and mixing. Both the fuel injector and mixer are located on a combustor dome plate or cap. 
     In general, the fuel to air ratio in the mixer is rich. Since the overall combustor fuel-air ratio of gas turbine combustors is lean, additional air is added through discrete dilution holes prior to exiting the combustor. Poor mixing and hot spots can occur both at the dome, where the injected fuel must vaporize and mix prior to burning, and in the vicinity of the dilution holes, where air is added to the rich dome mixture. In addition, many propulsion systems, such as those used in various tactical missile systems, involve an enclosed combustor. 
     Combustion instabilities are commonly encountered in low emissions gas turbine engines. Combustion dynamics in the form of fluctuations in pressure, heat-release rate, and other perturbations in flow may lead to problems such as structural vibration, excessive heat transfer to a chamber, and consequently lead to failure of the system. There are two basic methods for controlling combustion dynamics in a combustion system: passive control and active control. As the name suggests, passive control refers to a system that incorporates certain design features and characteristics to reduce dynamic pressure oscillations. Active control, on the other hand, incorporates a sensor or sensors to detect dynamics (e.g., pressure sensor to detect pressure fluctuations) and to provide a feedback signal which, when suitably processed by a controller, provides an input signal to a control device. The control device in turn operates to reduce the combustion instabilities. 
     The combustion characteristics of an enclosed combustor, including flammability limits, instability, and efficiency are closely related to the interaction between shear flow dynamics of the fuel and air flow at the inlet and acoustic modes of the combustor. Strong interaction, between the acoustic modes of the combustor and the airflow dynamics may lead to highly unstable combustion. Specifically, unstable combustion may occur when the acoustic modes of the combustor match the instability modes of the airflow. For such conditions, the shedding of the airflow vortices upstream of the combustor tends to excite acoustic resonances in the combustion chamber, which subsequently cause the shedding of more coherent energetic vortices at the resonant frequency. The continued presence of such vortices provides a substantial contribution to the instability of the combustion process. 
     In a jet of fluid that exits from a conduit to a surrounding medium of another fluid, sudden increase of the mass-flow leads to formation of well-defined vortices that dominate the boundary between the jet fluid and the surrounding fluid. Because these vortices help transport chunks of fluid over a large distance, the rate of turbulent mixing between the two fluids is closely linked to the dynamics of these vortices. One way to manipulate the dynamics of vortices is to modulate periodically the instantaneous mass-flux of the jet. 
     In combustion devices, actuators can be used to enhance combustion performance such as efficiency improvement, pollutant reduction, flammability extension, and instability suppression. Combustion apparatuses, which use actuators have been disclosed. One such disclosure includes several active control devices, including loudspeakers to modify the pressure field of the system or to obtain gaseous fuel flow modulations, pulsed gas jets aligned across a rearward facing step, adjustable inlets for time-variant change of the inlet area of a combustor, and solenoid-type fuel injectors for controlled unsteady addition of secondary fuel into the main combustion zone. 
     The periodic shedding of vortices produced in highly sheared gas flows has been recognized as a source of substantial acoustic energy for many years. For example, experimental studies have demonstrated that vortex shedding from gas flow restrictors disposed in large, segmented, solid propellant rocket motors couples with the combustion chamber acoustics to generate substantial acoustic pressures. The maximum acoustic energies are produced when the vortex shedding frequency matches one of the acoustic resonances of the combustor. It has been demonstrated that locating the restrictors near a velocity antinode generated the maximum acoustic pressures in a solid propellant rocket motor, with a highly sheared flow occurring at the grain transition boundary in boost/sustain type tactical solid propellant rocket motors. 
     Lean running engines tend to have flame extinguishment (also known as Lean Blow Out or LBO) or large pressure dynamics (known as combustion dynamics) inside the combustor that can be detrimental to engine operation and long term reliability. Lean premixed gas turbine combustors are prone to pressure fluctuations called combustion dynamics. Combustion dynamicsis a result of interaction between heat release from combusting fuel-air mixture and pressure oscillations in the combustion chamber. This phenomenon can result in expensive damage to combustor and or the gas turbine system hardware. 
     Therefore, there exists a need for a control system for combustors to operate near LBO boundaries without the risk of crossing the LBO boundary and also to near-simultaneously reduce combustion dynamics. 
     BRIEF DESCRIPTION OF THE INVENTION 
     In accordance with one embodiment of the present invention, an apparatus for active modulation of a flame in a combustor having instabilities, is provided. The apparatus includes an actuator, wherein the at least one actuator periodically pulses a fraction of the fuel flow delivered to the combustor. The apparatus also includes at least one sensor connected to the combustion chamber, wherein the sensor generates a signal in response to instabilities in the combustor. Lastly, the apparatus includes means for controlling the actuator in response to the sensor signals. The means for controlling prompts the actuator to periodically inject pulsed fuel into the combustor at a predetermined sympathetic frequency and magnitude, thereby controlling the amplitude of the oscillations in the combustor. 
     In accordance with another embodiment of the present invention, a method for active gas turbine combustion control of a combustor is provided. The method includes sensing a combustion dynamic signal and determining a sound pressure level associated with the combustion dynamic signal. The method performs a spectral analysis of the combustion dynamic signal and determines if the sound pressure level exceeds a predetermined threshold. The method then determines from the spectral analysis a plurality of sympathetic frequencies associated with the combustion dynamic signal and directs a MEMS actuator to inject pulsed fuel into the combustor in accordance with the sympathetic frequencies. 
     The invention is also directed towards an active combustion control (ACC) system for near simultaneously balancing lean blow out avoidance, combustor dynamics mitigation, and combustor operability. The ACC includes a combustor pressure sensor and a fuel modulating system for providing fuel to a combustor. The fuel modulation system further includes a MEMS microvalve for providing a predetermined fraction of the total combustion fuel flow upstream of at least one corresponding pilot fuel nozzle and at least one fuel flow sensor. The ACC also includes a combustion dynamics control system for receiving input from the combustor pressure sensors and providing combustion dynamics control signals to the fuel modulating system for modulating fuel to the combustor at a frequency and amplitude derived from determined sympathetic frequencies to reduce or cancel combustion dynamics. The ACC also includes a lean blow out (LBO) control system for receiving input from the combustor pressure sensors and providing a LBO control signal to the fuel modulating system for modulating fuel to the combustor at a frequency and amplitude derived from determined sympathetic frequencies to prevent lean blow out. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is an illustration of the high level architecture of the Active Combustion Control System utilizing fuel modulation system in accordance with an embodiment of the present invention; 
         FIG. 2  is an illustration of the high level architecture of the Active Combustion Control System in accordance with the invention shown in  FIG. 1 ; 
         FIG. 3  is an illustration of a combustor and control setup in accordance with the invention shown in  FIG. 1 ; 
         FIG. 4  is a cross-sectional view of a combustor used in accordance with the present invention shown in  FIG. 1 ; 
         FIG. 5  is a flow chart showing the behavior of the active combustion control system in accordance with the invention shown in  FIG. 1 ; 
         FIG. 6  is an illustration of an embodiment of MEMS for individual fuel nozzle flow trim/control.  FIG. 6  includes a pump  306 , fuel nozzle control  504 , and series of MEMs  202 . The system includes forty CPFN valves and twenty MEMs valves as shown in the  FIG. 6 . The other elements and sub-elements illustrated in  FIG.6  are self explanatory for explaining the various embodiments of the invention. 
         FIG. 7  is an illustration of another embodiment of MEMS for individual fuel nozzle trim/control.  FIG. 7  includes a pump  306 , fuel nozzle control  504 , and series of MEMs  202 . The system further includes forty CPFN valves and twenty MEMs valves as shown in the  FIG. 7 . The other elements and sub-elements illustrated in  FIG. 7  are self explanatory for explaining the various embodiments of the invention. 
         FIGS. 8A-8D  presents graphs demonstrating the effects of fuel forcing; 
         FIGS. 9A-9E  shows a spectrogram of the frequency response of the combustor  30  with no fuel forcing; and 
         FIGS. 10A-10C  shows data without and with fuel forcing. 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     Referring to  FIG. 1 , the ACCS employing fuel flow modulation system includes an upstream fuel metering &amp; pressurization system  200 , a split controller  201  for splitting fuel flow between cyclone valve  202   b  &amp; pilot valve  202   a , including suitable external inputs  201 A, a fuel modulation valve  202 , Valvevalve control electronics  203 , and a fuel flow sensor  204 . It will be appreciated that the fuel flow sensor  204  may be any type of suitable sensor such as a volumetric or mass fuel flow sensor. It will also be appreciated the sensor approaches may include piezoresistive cantilever, coriolis or anemometer approaches. Also shown in  FIG. 1  is a combustion diagnostic sensor  210 . Combustion diagnostic sensor  210  may include any suitable sensor such as a combustor dynamic pressure sensor, a combustor pressure sensor, a heat release sensor, an emissions sensor, and a Fuel to Air Ratio (F/A) measurement sensor located downstream of nozzle before flame. The system shown in  FIG. 1  includes an anti-aliasing/signal conditioning device  215 , a real time control platform  216 , and a signal conditioning device  217 . 
     The ACCS in  FIG. 1  controls fuel flow to a combustor  30 . The anti-aliasing/signal conditioning device accounts for sensor linearity, bandwidth, output (V or mA, etc.), and delay. The real time control platform  216  uses algorithms to control combustion dynamics, LBO, and emissions and provides an electrical signal to the signal conditioning device  217 . The signal conditioning device  217  conditions the signal from the real time control platform  216  accounting for bandwidth, linearity, amplification, etc. and sends the appropriate voltage/current requirements (not pressure actuated) to the valve control electronics  203 . 
     Referring also to  FIG. 2 , the Active Combustion Control System works by detecting lean blowout precursors and combustion dynamics within the combustion chamber  30  through pressure measurements from a combustion diagnostic sensor  210  such as a pressure sensor. In response to pressure oscillations typical of the onset of LBO or unacceptable combustion dynamics, amplitude and frequency modulations of fuel flow through nozzle ( FIG. 3 , item  162 ) is made in accordance with the teachings herein to prevent LBO or combustion dynamics. The modulation of the fuel flow is made by a fuel modulation valve  202  on the pilot circuit of the nozzle. Fuel modulation valve  202  may include Micro-Electro-Mechanical Systems (MEMS)-based microvalves or macroscale valves. Flow rate modulations are measured with a downstream fuel flow sensor  204  and high temperature flow sensor readout electronics  203   a  in order to determine a signal within the high temperature environment. The control system simultaneously balances lean blow out avoidance, combustor dynamics mitigation and combustor operability. 
     The fuel modulation valve  202  and the flow sensor  204  are located in the nozzle assembly while the diagnostic combustion sensor  210  is in the combustor  30 . In a particular embodiment, the fuel modulation valve  202  is able to actuate at a 1 kHz frequency at an elevated temperature and is able to modulate a significant portion of the pilot flow. The flow sensor  204  has a bandwidth capable of detecting the flow modulations of the fuel modulation valve  202 . The flow sensor readout electronics  203   a  associated with the flow sensor  204  are capable of measurements at elevated temperatures of about 200 C. so that it can be co-located with the flow sensor  204 . In a particular embodiment, the combustion diagnostic sensor  210  has a bandwidth greater than 1 kHz in order to detect combustion instabilities in the combustor  30  while having the capability to sustain reliable operation in a harsh environment. Using the combustion diagnostic sensor  210  response along with control algorithms, the fuel modulation valve  202  will modulate the flow into the combustor  30 . 
     The ACCS with fuel modulation system enables lower emission engines with less instability. It uses valve  202  that can operate at higher temperatures and at higher frequencies than conventional valves. The ACCS with fuel modulation system puts the small size flow sensor  204  in the system so that the flow pulsations and split levels can be monitored. 
     Referring to  FIG. 2 , there is shown a high level control architecture of the Active Combustion Control System (ACCS) in accordance with an embodiment of the present invention. The ACCS includes an engine controller  10 , engine control sensors  11 , combustion pressure dynamic sensors  12 , a LBO Precursor Detection  13 , a LBO Threshold Comparator  14 , a LBO Controller  15 , a Dynamics Mitigation Controller  16  and a Fuel Control System  17 . 
     In  FIG. 2 , the ACCS controls an Engine  18  and Combustor  30  (See  FIG. 1 ). The engine controller  10  modulates engine actuators (guide vanes, variable stator vanes, etc.) to achieve required performance while protecting engine against various constraints. The engine controller  10  also adjusts bulk fuel flow to meet engine performance requirements and pilot/cyclone fuel split for emissions. 
     One embodiment of the control system performs a real time spectral analysis of the dynamic pressure signal  12  from the combustor pressure sensor to determine the main pressure frequency in the combustor  30 , along with two minor sub-frequencies known as sympathetic frequencies (see  FIGS. 8-10 ). In general, the main frequency is the sum of the two sub-frequencies. The control system then sends a command to the fuel modulator system enabling that system to inject a fraction of the total combustion fuel flow upstream of the fuel nozzle at a predetermined frequency and amplitude. This is known as fuel forcing. 
     The modulated frequency coincides with either the peak or adjacent valley of one of the sympathetic frequencies. The fuel modulation at the lower frequency adds energy at the lower frequency with the result of shifting the main combustor frequency towards the lower frequency. The net result is the cancellation or at the minimum, a reduction in the sound pressure level of the observed combustion dynamics. Fuel forcing is maintained, with changes made to the frequency and amplitude until the undesirable engine dynamics are canceled, or until an acceptable pre-determined minimum dynamics level is attained. 
     Referring to  FIG. 2 , there is shown an illustration of the high level architecture of the Active Combustion Control System in accordance with the invention shown in  FIG. 1 ; In particular embodiments, the LBO controller  15  adjusts pilot/cyclone split via a MEMS valve  202  (refer to  FIG. 3 ) or bulk fuel flow. The dynamics mitigation controller  16  modulates MEMS valve  202  to introduce Fuel/Air (F/A) ratio (hence heat release) perturbation to mitigate dynamics. The LBO precursor detection  13  employs signal processing and data fusion algorithms. 
     Referring to  FIG. 3 , there is shown a pictorial diagram of one embodiment of the present invention. Fuel pump  306  provides fuel through fuel flow device  308  to fuel actuator  301 . Static pressure device  308 A measures the static pressure of the fuel prior to the fuel entering fuel actuator  301  and pilot injector fuel line  162 . A controller  304  controls fuel actuator  301  and also receives data from dynamic pressure sensor  308 B. The fuel actuator  301  is able to control the frequency and amplitude of the fuel in the pilot injector fuel line  162 . Pilot injector fuel line  162  is connected to combustor  30 . 
       FIG. 4  is a cross-sectional view of a combustor  30  for use with a gas turbine engine. In one embodiment, the gas turbine engine is a GE F414 engine available from General Electric Company, Cincinnati, Ohio. This example is illustrative, and the invention is not limited to specific engine models or combustor designs. In the illustrated example, combustor  30  includes an annular outer liner  40 , an annular inner liner  42 , and a domed inlet end  44  extending between outer and inner liners  40  and  42 , respectively. Domed inlet end  44  has a shape of a low area ratio diffuser. 
     Outer liner  40  and inner liner  42  are spaced radially inward from a combustor casing  46  and define a combustion chamber  48 . Combustor casing  46  is generally annular and extends downstream from an exit  50  of a compressor, such-as compressor  14  shown in  FIG. 4 . Combustion chamber  48  is generally annular in shape and is disposed radially inward from liners  40  and  42 . Outer liner  40  and combustor casing  46  define an outer passageway  52  and inner liner  42  and combustor casing  46  define an inner passageway  54 . Outer and inner liners  40  and  42 , respectively, extend to a turbine inlet nozzle  58  disposed downstream from diffuser  48 . 
     A trapped vortex cavity  70  is incorporated into a portion  72  of outer liner  40  immediately downstream of dome inlet end  44 . Trapped vortex cavity  70  has a rectangular cross-sectional profile and because trapped vortex cavity  70  opens into combustion chamber  48 , cavity  70  only includes an aft wall  74 , an upstream wall  76 , and an outer wall  78  extending between aft wall  74  and upstream wall  76 . In an alternative embodiment, trapped vortex cavity  70  has a non-rectangular cross-sectional profile. In a further alternative embodiment, trapped vortex cavity  70  includes rounded comers. Outer wall  78  is substantially parallel to outer liner  40  and is radially outward a distance  80  from outer liner  40 . A corner bracket  82  extends between trapped vortex cavity aft wall  74  and combustor outer liner  40  and secures aft wall  74  to outer liner  40 . Trapped vortex cavity upstream wall  76 , aft wall  74 , and outer wall  78  each include a plurality of passages (not shown) and openings (not shown) to permit air to enter trapped vortex cavity  70 . 
     Trapped vortex cavity upstream wall  76  also includes an opening  86  sized to receive a fuel injector assembly  90 . Fuel injector assembly  90  extends radially inward through combustor casing  46  upstream from a combustion chamber upstream wall  92  defining combustion chamber  48 . Combustion chamber upstream wall  92  extends between combustor inner liner  42  and trapped vortex cavity upstream wall  76  and includes an opening  94 . Combustion chamber upstream wall  92  is substantially co-planar with trapped vortex cavity upstream wall  76 , and substantially perpendicular to combustor inner liner  42 . 
     Combustor upstream wall opening  94  is sized to receive a mixer assembly  96 . Mixer assembly  96  is attached to combustion chamber upstream wall  92  such that a mixer assembly axis of symmetry  98  is substantially co-axial with an axis of symmetry  99  for combustion chamber  48 . Mixer assembly  96  is generally cylindrical-shaped with an annular cross-sectional profile (not shown) and includes an outer wall  100  that includes an upstream portion  102  and a downstream portion  104 . 
     Mixer assembly outer wall upstream portion  102  is substantially cylindrical and has a diameter  106  sized to receive fuel injector assembly  90 . Mixer assembly outer wall downstream portion  104  extends from upstream portion  102  to combustor upstream wall opening  94  and converges towards mixer assembly axis of symmetry  98 . Accordingly, a diameter  110  of upstream wall opening  94  is less than upstream portion diameter  106 . 
     Mixer assembly  96  also includes a swirler  112  extending circumferentially within mixer assembly  96 . Swirler  112  includes an intake side  114  and an outlet side  116 . Swirler  112  is positioned adjacent an inner surface  118  of mixer assembly outer wall upstream portion  102  such that swirler intake side  114  is substantially co-planar with a leading edge &#39; 120  of mixer assembly outer wall upstream portion  102 . Swirler  112  has an inner diameter  122  sized to receive fuel injector assembly  90 . In one embodiment, swirlers  112  are single axial swirlers. In an alternative embodiment, swirlers  112  are radial swirlers. 
     Fuel injector assembly  90  extends radially inward into combustor  16  through an opening  130  in combustor casing  46 . Fuel injector assembly  90  is positioned between domed inlet end  44  and mixer assembly  96  and includes a pilot fuel injector  140  and a main fuel injector  142 . Main fuel injector  142  is radially inward from pilot fuel injector  140  and is positioned within mixer assembly  96  such that a main fuel injector axis of symmetry  144  is substantially co-axial with mixer assembly axis of symmetry  98 . Specifically, main fuel injector  142  is positioned such that an intake side  146  of main fuel injector  142  is upstream from mixer assembly  96  and a trailing end  148  of main fuel injector  142  extends through mixer assembly  96  radially inward from swirler  112  and towards combustor upstream wall opening  94 . Accordingly, main fuel injector  142  has a diameter  150  that is slightly less than swirler inner diameter  122 . 
     Pilot fuel injector  140  is radially outward from main fuel injector  142  and is positioned upstream from trapped vortex cavity upstream wall opening  86 . Specifically, pilot fuel injector  140  is positioned such that a trailing end  154  of pilot fuel injector  140  is in close proximity to opening  86 . 
     A fuel delivery system  160  supplies fuel to combustor  30  and includes a pilot fuel circuit  162  and a main fuel circuit  164  to control nitrous oxide emissions generated within combustor  30 . Pilot fuel circuit  162  supplies fuel to trapped vortex cavity  70  through fuel injector assembly  90  and main fuel circuit  164  supplies fuel to mixer assembly  96  through fuel injector assembly  90 . During operation, as gas turbine engine  10  is started and operated at idle operating conditions, fuel and air are supplied to combustor  30 . During gas turbine idle operating conditions, combustor  30  uses only the pilot fuel stage for operating. Pilot fuel circuit  162  injects fuel to combustor trapped vortex cavity  70  through pilot fuel injector  140 . Simultaneously, airflow enters trapped vortex cavity  70  through aft, upstream, and outer wall air passages and enters mixer assembly  96  through swirlers  112 . The trapped vortex cavity air passages form a collective sheet of air that mixes rapidly with the fuel injected and prevents the fuel from forming a boundary layer along aft wall  74 , upstream wall  76 , or outer wall  78 . 
     Combustion gases  180  generated within trapped vortex cavity  70  swirl in a counter-clockwise motion and provide a continuous ignition and stabilization source for the fuel/air mixture entering combustion chamber  48 . Airflow  182  entering combustion chamber  48  through mixer assembly swirler  112  increases a rate of fuel/air mixing to enable substantially near-stoichiometric flame-zones (not shown) to propagate with short residence times within combustion chamber  48 . 
     Referring to  FIG. 5 , there is shown a flow chart showing the behavior of the active combustion control system in accordance with the invention shown in  FIG. 1 . Sensors within or located substantially near combustor ( FIG. 4 , item  30 ) determine combustor sound pressure level magnitude associated with a combustion dynamic signal  1101 . 
     A threshold comparator ( FIG. 2 . item  14 A) determines  1102  if the magnitude is below a predetermined threshold. If the threshold comparator ( FIG. 2 . item  14 A) determines the magnitude is below the predetermined threshold, control is passed back to the sensors for determining the combustor dynamic signal  1101 . 
     If the threshold comparator ( FIG. 2 . item  14 A) determines the magnitude is above the predetermined threshold control is passed to a spectral analysis module co-located within the lean-blow-out detector ( FIG. 2 , item  13 A) for performing  1103  a spectral analysis of the combustor dynamic signal. 
     In addition an acoustic analysis module co-located within the lean-blow-out precursor detector ( FIG. 2 , item  13 B) also performs  1104  an acoustic analysis of the combustor dynamic signal. 
     The lean-blow-out (LBO) precursor detector ( FIG. 2 , item  15 ) determines  1105  sympathetic frequencies associated with a dominant instability frequency of the combustor dynamic signal. It will be appreciated that the sympathetic frequencies are generally much lower frequency than the dominant instability frequency. 
     LBO controller ( FIG. 2 , item  15 ) determines  1107  an optimum fuel forcing frequency and magnitude based upon the sympathetic frequencies. LBO controller then sends  1108  command signals to MEMS fuel actuator ( FIG. 1 , item  202 ) to force fuel input to the combustor ( FIG. 4 , item  30 ) pilot input port ( FIG. 4 , item  162 ) at the determined optimum fuel forcing frequency and magnitude. 
     Experimental Results 
     Referring also to  FIG. 3  there is shown a depiction of a test setup with a swirl stabilized combustor  30 , a premixing fuel injector  308 , a high-speed fuel actuator  301  and a controller  304 . This setup is representative of that used on the tunable acoustic test rig (TCA) tests. 
       FIGS. 8A-8E  show data from the TCA single nozzle test rig (e.g.,  FIG. 3 ) operating with a twin annular premixing swirler (TAPS) premixing fuel injector. The tests were run with liquid Jet-A fuel. 
       FIGS. 9A-9E  shows a spectrogram of the frequency response of the combustor  30  with no fuel forcing. Shifts in the dominant frequencies that show up on the spectrogram are due to the changing length of the combustor section. Analysis of unforced TCA rig data shows fundamental, its harmonics (in red) and various sums and differences (in green) of these frequencies 
       FIGS. 10A-10C  shows data without and with fuel forcing. KP represents dynamic pressure signals, while valve shows the fuel actuator frequency command. A reduction in pressure fluctuation amplitude is shown from about 45 s when the fuel is actuated with a square wave at 143 Hz. This suppression is shown in the time domain from 45-75 s, after which the signal strength begins to build up again after the fuel forcing stopped. 
     This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to make and use the invention. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.