Abstract:
A fuel nozzle for a turbine engine has a central body member with a pilot, a surrounding barrel housing, a mixing duct and an air inlet duct. The fuel nozzle additionally has a main fuel injection device located between the air inlet duct and the mixing duct. The main fuel injection device is configured to introduce a flow of fuel into the barrel member to create a fuel/air mixture which is then premixed with a swirler. The fuel/air mixture then further mixes in the mixing duct and exits the nozzle into a combustor for combustion. The geometry of the fuel nozzle ensures that pressure waves from the combustor do not create a time varying fuel to air equivalence ratio in the flow through the nozzle that achieves a resonance with the pressure waves.

Description:
This application is a continuation of U.S. patent application Ser. No. 12/841,140 filed Jul. 21, 2010, which is a divisional of U.S. patent application Ser. No. 11/239,376, filed Sep. 30, 2005, now abandoned. 
    
    
     TECHNICAL FIELD 
     The present disclosure relates generally to a turbine engine, and more particularly, to a turbine engine having an acoustically tuned fuel nozzle. 
     BACKGROUND 
     Internal combustion engines, including diesel engines, gaseous-fueled engines, and other engines known in the art, may exhaust a complex mixture of air pollutants. These air pollutants may be composed of gaseous compounds, which may include nitrous oxides (NOx). Due to increased attention on the environment, exhaust emission standards have become more stringent and the amount of NOx emitted to the atmosphere from an engine may be regulated depending on the type of engine, size of engine, and/or class of engine. 
     It has been established that a well-distributed flame having a low flame temperature can reduce NOx production to levels compliant with current emission regulations. One way to generate a well-distributed flame with a low flame temperature is to premix fuel and air to a predetermined lean fuel to air equivalence ratio. However, naturally-occurring pressure fluctuations within the turbine engine can be amplified during operation of the engine under these lean conditions. In fact, the amplification can be so severe that damage and/or failure of the turbine engine can occur. 
     One method that has been implemented by turbine engine manufacturers to provide lean fuel/air operational conditions within a turbine engine while minimizing the harmful vibrations generally associated with lean operation is described in U.S. Pat. No. 6,698,206 (the &#39;206 patent) issued to Scarinci et al. on Mar. 2, 2004. The &#39;206 patent describes a turbine engine having a primary combustion zone, a secondary combustion zone, and a tertiary combustion zone. Each of the combustion zones is supplied with premixed fuel and air by respective mixing ducts and a plurality of axially spaced-apart air injection apertures. These apertures reduce the magnitude of fluctuations in the lean fuel to air equivalence ratio of the fuel and air mixtures supplied into the mixing zones, thereby reducing the harmful vibrations. 
     Although the method described in the &#39;206 patent may reduce some harmful vibrations associated with a low NOx-emitting turbine engine, it may be expensive and insufficient. In particular, the many apertures associated with each of the combustion zones described in the &#39;206 patent may drive up the cost of the turbine engine. In addition, because the reduction of vibration within the turbine engine of the &#39;206 patent does not rely upon strategic placement of the apertures according to acoustic tuning specific to the particular turbine engine, the reduction of vibration may be limited and, in some situations, insufficient. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a cutaway-view illustration of an exemplary disclosed turbine engine; 
         FIG. 2  is a cross-sectional illustration of an exemplary disclosed fuel nozzle for the turbine engine of  FIG. 1 ; and 
         FIG. 3  is a pictorial representation of an exemplary disclosed operation of the fuel nozzle of  FIG. 2 . 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  illustrates an exemplary turbine engine  10 . Turbine engine  10  may be associated with a stationary or mobile work machine configured to accomplish a predetermined task. For example, turbine engine  10  may embody the primary power source of a generator set that produces an electrical power output or of a pumping mechanism that performs a fluid pumping operation. Turbine engine  10  may alternatively embody the prime mover of an earth-moving machine, a passenger vehicle, a marine vessel, or any other mobile machine known in the art. Turbine engine  10  may include a compressor section  12 , a combustor section  14 , a turbine section  16 , and an exhaust section  18 . 
     Compressor section  12  may include components rotatable to compress inlet air. Specifically, compressor section  12  may include a series of rotatable compressor blades  22  fixedly connected about a central shaft  24 . As central shaft  24  is rotated, compressor blades  22  may draw air into turbine engine  10  and pressurize the air. This pressurized air may then be directed toward combustor section  14  for mixture with a liquid and/or gaseous fuel. It is contemplated that compressor section  12  may further include compressor blades (not shown) that are separate from central shaft  24  and remain stationary during operation of turbine engine  10 . 
     Combustor section  14  may mix fuel with the compressed air from compressor section  12  and combust the mixture to create a mechanical work output. Specifically, combustor section  14  may include a plurality of fuel nozzles  26  annularly arranged about central shaft  24 , and an annular combustion chamber  28  associated with fuel nozzles  26 . Each fuel nozzle  26  may inject one or both of liquid and gaseous fuel into the flow of compressed air from compressor section  12  for ignition within combustion chamber  28 . As the fuel/air mixture combusts, the heated molecules may expand and move at high speed into turbine section  16 . 
     As illustrated in the cross-section of  FIG. 2 , each fuel nozzle  26  may include components that cooperate to inject gaseous and liquid fuel into combustion chamber  28 . Specifically, each fuel nozzle  26  may include a barrel housing  34  connected on one end to an air inlet duct  35  for receiving compressed air, and on the opposing end to a mixing duct  37  for communication of the fuel/air mixture with combustion chamber  28 . Fuel nozzle  26  may also include a central body  36  with a pilot fuel injector, and a swirler  40 . Central body  36  may be disposed radially inward of barrel housing  34  and aligned along a common axis  42 . A pilot fuel injector may be located within central body  36  and configured to inject a pilot stream of pressurized fuel through a tip end  44  of central body  36  into combustion chamber  28  to facilitate engine starting, idling, cold operation, and/or lean burn operations of turbine engine  10 . Swirler  40  may be annularly disposed between barrel housing  34  and central body  36 . 
     Barrel housing  34  may embody a tubular member having a plurality of air jets  46 . Air jets  46  may be co-aligned at a predetermined axial position along the length of barrel housing  34 . This predetermined axial position may be set during manufacture of turbine engine  10  to attenuate a time-varying flow of air entering fuel nozzle  26  via air inlet duct  35 . It is contemplated that air jets  46  may be located at any axial position along the length of barrel housing  34  and may vary from engine to engine or from one class or size of engine to another class or size of engine according to attenuation requirements. Air jets  46  may receive compressed air from compressor section  12  by way of one or more fluid passageways (not shown) external to barrel housing  34 . 
     Air inlet duct  35  may embody a tubular member configured to axially direct compressed air from compressor section  12  (referring to  FIG. 1 ) to barrel housing  34 , and to divert a portion of the compressed air to air jets  46 . Specifically, air inlet duct  35  may include a central opening  48  and a flow restrictor  50  located within central opening  48  at an end opposite barrel housing  34 . In one example, flow restrictor  50  may embody a blocker ring extending inward from the interior surface of air inlet duct  35 . The radial distance that flow restrictor  50  protrudes into central opening  48  may determine the amount of compressed air diverted around air inlet duct  35  to air jets  46  during operation of turbine engine  10 . The amount of air diverted to air jets  46  may be less than the amount of air passing through air inlet duct  35 . The geometry of air inlet duct  35  may such that pressure fluctuations within fuel nozzle  26  may be minimized to provide for piece-wise uniform flow through air inlet duct  35 . In one example, air inlet duct  35  may be generally straight and may have a predetermined length. The predetermined length of air inlet duct  35  may be set during manufacture of turbine engine  10  according to an axial fuel introduction location and a naturally-occurring pressure fluctuation with combustion chamber  28 . The method of determining and setting the length of air inlet duct  35  will be discussed in more detail below. 
     Mixing duct  37  may embody a tubular member configured to axially direct the fuel/air mixture from fuel nozzle  26  into combustion chamber  28 . In particular, mixing duct  37  may include a central opening  52  that fluidly communicates barrel housing  34  with combustion chamber  28 . The geometry of mixing duct  37  may be such that pressure fluctuations within fuel nozzle  26  are minimized to provide for piece-wise uniform flow through air inlet duct  35 . In one example, mixing duct  37  may be generally straight and may have a predetermined length. Similar to air inlet duct  35 , the predetermined length of mixing duct  37  may be set during manufacture of turbine engine  10  according to an axial fuel introduction location and the naturally-occurring pressure fluctuation within combustion chamber  28 . The method of determining and setting the length of mixing duct  37  will be discussed in more detail below. 
     Swirler  40  may be situated to radially redirect an axial flow of compressed air from air inlet duct  35 . In particular, swirler  40  may embody an annulus having a plurality of connected vanes  54  located within an axial flow path of the compressed air. As the compressed air contacts vanes  54 , it may be diverted in a radially inward direction. It is contemplated that vanes  54  may extend from barrel housing  34  radially inward directly toward common axis  42  or, alternatively, to a point centered off-center from common axis  42 . It is also contemplated that vanes  54  may be straight or twisted along a length direction and tilted at an angle relative to an axial direction of common axis  42 . 
     Vanes  54  may facilitate fuel injection within barrel housing  34 . In particular, some or all of vanes  54  may each include a liquid fuel jet  56  and a plurality of gaseous fuel jets  58 . It is contemplated that any number or configuration of vanes  54  may include liquid fuel jets  56 . The location of vanes  54  along common axis  42  and the resulting axial fuel introduction point within fuel nozzle  26  may vary and be set to, in combination with specific time-varying air flow characteristics, attenuate the naturally-occurring pressure fluctuation within combustion chamber  28 . The method of determining and setting the axial fuel introduction point will be discussed in more detail below. 
     Gaseous fuel jets  58  may provide a substantially constant mass flow of gaseous fuel such as, for example, natural gas, landfill gas, bio-gas, or any other suitable gaseous fuel to combustion chamber  28 . In particular, gaseous fuel jets  58  may embody restrictive orifices (i.e., gaseous fuel jets  58  may include an exit port comprising a restriction to the fuel exiting into barrel housing  34 ), situated along a leading edge of each vane  54 . Each of gaseous fuel jets  58  may be in communication with a central fuel passageway  59  within the associated vane  54  to receive gaseous fuel from an external source (not shown). The restriction, i.e., exit port, at gaseous fuel jets  58  may be the greatest restriction applied to the flow of gaseous fuel within fuel nozzle  26 , such that a substantially continuous mass flow of gaseous fuel from gaseous fuel jets  58  may be ensured. 
     Combustion chamber  28  (referring to  FIG. 1 ) may house the combustion process. In particular, combustion chamber  28  may be in fluid communication with each fuel nozzle  26  and may be configured to receive a substantially homogenous mixture of fuel and compressed air. The fuel/air mixture may be ignited and may fully combust within combustion chamber  28 . As the fuel/air mixture combusts, hot expanding gases may exit combustion chamber  28  and enter turbine section  16 . 
     Turbine section  16  may include components rotatable in response to the flow of expanding exhaust gases from combustor section  14 . In particular, turbine section  16  may include a series of rotatable turbine rotor blades  30  fixedly connected to central shaft  24 . As turbine rotor blades  30  are bombarded with high-energy molecules from combustor section  14 , the expanding molecules may cause central shaft  24  to rotate, thereby converting combustion energy into useful rotational power. This rotational power may then be drawn from turbine engine  10  and used for a variety of purposes. In addition to powering various external devices, the rotation of turbine rotor blades  30  and central shaft  24  may drive the rotation of compressor blades  22 . 
     Exhaust section  18  may direct the spent exhaust from combustor and turbine sections  14 ,  16  to the atmosphere. It is contemplated that exhaust section  18  may include one or more treatment devices configured to remove pollutants from the exhaust and/or attenuation devices configured to reduce the noise associated with turbine engine  10 , if desired. 
       FIG. 3  illustrates an exemplary relationship between the length of air inlet duct  35 , the length of mixing duct  37 , the axial fuel introduction point within barrel housing  34  resulting from the position of swirler  40  along common axis  42 , and the naturally-occurring pressure fluctuation stemming from a flame front  67  within combustion chamber  28 .  FIG. 3  will be discussed in more detail below. 
     INDUSTRIAL APPLICABILITY 
     The disclosed fuel nozzle may be applicable to any turbine engine where reduced vibrations within the turbine engine are desired. Although particularly useful for low NOx-emitting engines, the disclosed fuel nozzle may be applicable to any turbine engine regardless of the emission output of the engine. The disclosed fuel nozzle may reduce vibrations by acoustically attenuating a naturally-occurring pressure fluctuation within a combustion chamber of the turbine engine. The operation of fuel nozzle  26  will now be explained. 
     During operation of turbine engine  10 , air may be drawn into turbine engine  10  and compressed via compressor section  12  (referring to  FIG. 1 ). This compressed air may then be axially directed into combustor section  14  and against vanes  54  of swirler  40 , where the flow may be redirected radially inward. As the flow of compressed air is turned to flow radially inward, liquid fuel may be injected from liquid fuel jets  56  for mixing prior to combustion. Alternatively or additionally, gaseous fuel may be injected from gaseous fuel jets  58  for mixing with the compressed air prior to combustion. As the mixture of fuel and air enters combustion chamber  28 , it may ignite and fully combust. The hot expanding exhaust gases may then be expelled into turbine section  16 , where the molecular energy of the combustion gases may be converted to rotational energy of turbine rotor blades  30  and central shaft  24 . 
       FIG. 3  illustrates the time-varying flow characteristics of fuel and air entering fuel nozzle  26  and their effects on the naturally-occurring pressure fluctuations within combustion chamber  28 . In particular,  FIG. 3  illustrates a first curve  60 , a second curve  62 , a third curve  64 , and a plurality of pressure pulses  66 . First curve  60  may represent the time-varying flow of compressed air entering fuel nozzle  26  via air inlet duct  35 . Second curve  62  may represent the time-varying flow of fuel flow entering fuel nozzle  26  via liquid and/or gaseous fuel jets  56 ,  58 . Third curve  64  may represent the time-varying fuel to air equivalence ratio Φ (e.g., the instantaneous ratio of the amount of fuel within any axial plane along the length of fuel nozzle  26  to the amount of air in the same axial plane). Pressure pulses  66  may represent a wave of pressure traveling from combustion chamber  28  in a reverse direction toward air inlet duct  35  as a result of combustion within combustion chamber  28 . 
     Pressure pulses  66  may affect the time-varying characteristic of first, second, and third curves  60 - 64 . Specifically, as pressure pulses  66  travel in the reverse direction within fuel nozzle  26  and reach liquid and gaseous fuel injectors  56 ,  58  and the entrance to air inlet duct  35 , the pressure of each pulse may cause the flow rate of fuel and air entering fuel nozzle  26  to vary. These varying flow rates correspond to the amplitude variations of first and second curves  60 ,  62  illustrated in  FIG. 3 , which equate to the varying amplitude and phase angle of third curve  64 . When the value of Φ at the point of combustion within combustion chamber  28  is high compared to a time average value of Φ, the heat release and resulting pressure wave within combustion chamber  28  may be high. Likewise, when the value of Φ at the point of combustion within combustion chamber  28  is low compared to the time average value of Φ, the heat release and resulting pressure wave within combustion chamber  28  may be low. 
     Damage may occur when the phase angle of third curve  64  and the wave of pressure pulses  66  near alignment. That is, when the value of Φ entering combustion chamber  28  is high compared to the time average of Φ and enters combustion chamber  28  at about the same time that a pressure pulse  66  initiates from a flame front with combustion chamber  28 , resonance may be attained. Likewise, if the value of Φ entering combustion chamber  28  is low compared to the time average of Φ and enters combustion chamber  28  at a time between the initiation of pressure pulses  66 , resonance may be attained. It may be possible that this resonance could amplify pressure pulses  66  to a damaging magnitude. 
     Damage may be prevented when third curve  64  and the wave of pressure pulses  66  are out of phase. In particular, if the value of Φ entering combustion chamber  28  is low compared to the time average of Φ and enters combustion chamber  28  at the same time that a pressure pulse  66  initiates from a flame front within combustion chamber  28 , attenuation of pressure pulse  66  may be attained. Likewise, if the value of Φ entering combustion chamber  28  is high compared to the time average of Φ and enters combustion chamber  28  at a time between the initiation of pressure pulses  66 , attenuation may be attained. Attenuation could lower the magnitude of pressure pulses  66 , thereby minimizing the likelihood of damage to turbine engine  10 . 
     The phase angle and magnitude of Φ may be affected by the length of air inlet duct  35 , the length of mixing duct  37 , the axial fuel introduction point, and the axial location of air jets  46 . Specifically, by increasing the length of air inlet duct  35  (e.g., extending the entrance of air inlet duct  35  leftward, when viewed in  FIG. 2 ), the phase angle of first curve  60  may likewise shift to the left. In contrast, by decreasing the length of air inlet duct  35  (e.g., moving the entrance of air inlet duct  35  to the right, when viewed in  FIG. 2 ), the phase angle of first curve  60  may likewise move to the right. In fact, if the length of air inlet duct  35  becomes so short that the introduction of air is substantially coterminous with the introduction of fuel via gaseous fuel jets  58  and the pressure drops across flow restrictor  50  and gaseous fuel jets  58  are substantially constant, the phase angle and amplitude differences between first and second curves  60 ,  62  may be nearly zero, resulting in a substantially constant value of Φ. In addition, by extending the length of mixing duct  37  (e.g., extending the exit of mixing duct  37  rightward, when viewed  FIG. 2 ), the phase angle of first curve  60  may move to the left. By decreasing the length of mixing duct  37  (e.g., moving the exit of mixing duct  37  leftward, when viewed in  FIG. 2 ), the phase angle of first curve  60  may move to the right. By moving the location of swirler  40  left or right and, in doing so, the axial introduction point of gaseous and liquid fuel left or right, the phase angle of second curve  62  may mimic the same shifts. As the phase angle of one or both of first and second curves  60 ,  62  shifts, the phase angle and amplitude of third curve  64  may be affected. In this manner, the value of Φ entering combustion chamber  28  can be acoustically tuned to attenuate the naturally-occurring pressure pulses  66  of a specific engine or specific class or size of engine. It is contemplated that only one or both of the lengths of air inlet duct  35  and mixing duct  37  may be modified to attenuate the naturally-occurring pressure pulses  66 . 
     Further reduction in the magnitude of pressure pulses  66  may be attained by providing a substantially time-constant value of Φ. One way to reduce the variation in the value of Φ may be to reduce the time-varying characteristic of first and/or second curves  60 ,  62 . The time-varying characteristic of gaseous fuel introduced into combustion chamber  28  via gaseous fuel jets  58  may be reduced by way of the restriction at the surface of gaseous fuel jets  58 . This restriction may increase the pressure drop across gaseous fuel jets  58  to a magnitude at which the pressure fluctuations within fuel nozzle  26  may have little affect on the flow of fuel through gaseous fuel jets  58 . Another way to reduce the vibrations may be realized through the use of air jets  46 . In particular, as seen in  FIG. 3 , when pulses of compressed air are introduced at a specific location within fuel nozzle  26  and at a timing out of phase with first curve  60 , the time-varying characteristic of air entering combustion chamber  28  may be attenuated. In one example, the pulses of compressed air may be injected by air jets  46  substantially 180 degrees out of phase with first curve  60 . The affect of the injected pulses of air can be seen in  FIG. 3 ; as the flow of compressed air entering barrel housing  34  via air inlet duct  35  passes in proximity to air jets  46 , the amplitude of first curve  60  may be reduced. 
     Several advantages over the prior art may be associated with fuel nozzle  26  of turbine engine  10 . Specifically, because the length of air inlet duct  35 , the length of mixing duct  37 , and the axial fuel introduction point of turbine engine  10  may be selected specifically to attenuate the naturally-occurring pressure pulses of combustion chamber  28 , harmful vibrations of turbine engine  10  may be greatly reduced. This acoustic tuning of turbine engine  10  may be more successful at reducing vibration than the random placement of apertures in an attempt to create non-resonating turbulence. In addition, these reductions in vibration may be attained with minimal changes to existing hardware, resulting in lower component costs of turbine engine  10 . 
     It will be apparent to those skilled in the art that various modifications and variations can be made to the disclosed fuel nozzle. Other embodiments will be apparent to those skilled in the art from consideration of the specification and practice of the disclosed fuel nozzle. It is intended that the specification and examples be considered as exemplary only, with a true scope being indicated by the following claims and their equivalents.