Patent Application
20140033719
The present application and the resultant patent relate to gas turbine engines and more particularly relate to a multi-step combustor with a number of pilot fuel/air lines so as to mitigate vortex driven combustion instabilities for reduced emission levels and increased power output.
Operational efficiency and output of a gas turbine engine increases with increases in the temperature of the hot combustion gases. High combustion gas temperatures, however, may produce high levels of nitrogen oxides (NOx) and other types of regulated emissions. A balancing act thus exists between operating a gas turbine engine in an efficient temperature range while also ensuring that the output of nitrogen oxides and other types of regulated emissions remain below mandated levels.
Lean premixing tends to reduce combustion temperatures and the output of nitrogen oxides. A gas turbine engine thus may be operated in a lean premixed regime to achieve lower emission levels of nitrogen oxides. Lean premixed combustors, however, may be more susceptible to combustion instabilities due to pressure oscillations in the combustion chamber. Such instabilities may cause undesirable acoustic noise, reduce engine performance and reliability, and/or increase the frequency of required service. Flow coherent structures may play a critical role in driving low frequency combustion instabilities. The flow structures or vortices may be formed by the interaction between sheer flow instabilities and the acoustic resonance of the chamber. When these vortices dominate the reacting flow, the coherent flow structures may lead to periodic heat release and may result in high amplitude pressure oscillations. High levels of combustion dynamics thus may limit the operability envelope of the combustor in terms of emissions and/or power output.
There is thus a need for an improved combustor design so as to limit the impact of low frequency combustion instabilities. Mitigating and controlling such dynamics should improve overall mixing and flame stability for improved emissions and power output. Moreover, the operating life of the combustor and the overall gas turbine may be improved.
The present application and the resultant patent thus provide a combustor for use with a gas turbine engine. The combustor may include a number of fuel nozzles and a combustion zone downstream of the fuel nozzles. The combustion zone may include a number of steps such that the combustion zone expands in a radial direction downstream of the fuel nozzles.
The present application and the resultant patent further provide a method of limiting combustion instabilities in a combustor. The method may include the steps of introducing a fuel/air mixture into a multi-step combustion zone, introducing a pilot fuel/air mixture into the multi-step combustion zone, and altering an equivalence ratio of the fuel/air mixture.
The present application and the resultant patent further provide a combustor for use with a gas turbine engine. The combustor may include a number of fuel nozzles and a combustion zone downstream of the fuel nozzles. The combustion zone may include a number of steps such that the combustion zone expands in a radial direction and a number of pilot fuel/air lines therein.
These and other features and improvements of the present application and the resultant patent will become apparent to one of ordinary skill in the art upon review of the following detailed description when taken in conjunction with the several drawings and the appended claims.
Referring now to the drawings, in which like numerals refer to like elements throughout the several views,
The gas turbine engine 10 may use natural gas, various types of syngas, and/or other types of fuels. The gas turbine engine 10 may be any one of a number of different gas turbine engines offered by General Electric Company of Schenectady, New York, including, but not limited to, those such as a 7 or a 9 series heavy duty gas turbine engine and the like. The gas turbine engine 10 may have different configurations and may use other types of components. Other types of gas turbine engines also may be used herein. Multiple gas turbine engines, other types of turbines, and other types of power generation equipment also may be used herein together.
The combustor 100 also may include a number of pilot fuel/air lines 160. The pilot fuel/air lines 160 may be positioned about the combustion zone 120 so as to inject a pilot fuel/air mixture 170 in a largely radial direction 180, i.e., perpendicular to the direction of the fuel/air mixture 90. The number of the pilot fuel/air lines 160 may vary. The size, shape, and configuration of the pilot fuel/air lines 160 also may vary. The nature of the pilot fuel/air mixture 170 may vary. Each step 140 may have one or more of the pilot fuel/air lines 160 therein. Other components and other configurations also may be used herein.
In use, the multi-step combustion zone 130 may have a direct impact on the length scale of the vortices 190 and the shedding frequencies within the combustion zone 130. In other words, for a given Strouhal number, the reduced height of the steps 140 increases the shedding frequency and hence turbulence so as to minimize vortex driven combustion instabilities therein. For example, the shear layer may be separated from an upstream step so as to impinge on the next downstream step edge and, hence, act as a source of turbulence production. Such increased turbulence may prevent the development of large scale structures in the flow so as to enhance fine scale mixing and flame stability. The use of the steps 140 also has an impact on the time delay in heat release fluctuations and other advantages.
The use of the pilot fuel/air lines 160 about the combustion zone 130 helps in controlling the flow flame-acoustic interaction. Specifically, the pilot fuel/air mixture 170 may have an impact on the local equivalence ratio of the fuel/air mixture 90 at each step 140. The phase between the heat release and the acoustic pressure thus may play a role in feeding the energy into the acoustic nodes. This phase modification may be achieved by altering the equivalence ratio by injecting the pilot fuel/air mixture 170 into the fuel/air mixture 90. The addition of the pilot fuel/air mixture 170 thus changes the heat release distribution therein. The injection of the pilot fuel/air mixture 170 may be adjusted such that the flame does not add energy into the acoustic nodes. Injection of the pilot fuel/air mixture 170 also affects the NOx emissions profile. Overall operation of the gas turbine 10 thus may be improved as well as the overall operating life.
The use of the steps 140 may be optimized herein based upon the size of the vortices and the vortex shedding frequency. The pilot fuel/air mixture 170 may be injected in either the radial direction 180 or the axial direction 210 so as to change the overall equivalence ratio. The combustors 100, 200 described herein thus provide both passive control through the use of the multi-step combustion zone 120 as well as active control given the pilot fuel/air lines 160. The combination of the passive and active controls thus may extend the range of stable operating conditions for the combustors 100, 200 herein and the like.
It should be apparent that the foregoing relates only to certain embodiments of the present application and the resultant patent. Numerous changes and modifications may be made herein by one of ordinary skill in the art without departing from the general spirit and scope of the invention as defined by the following claims and the equivalents thereof.
