Patent Application
20090139238
The present invention applies generally to gas turbine combustors and more specifically to an apparatus and method for providing improved combustion stability and lower pressure drop across the combustion system.
In a combustion system for a gas turbine, fuel and compressed air are mixed together and ignited to produce hot combustion gases that drive a turbine and produce thrust or drive a shaft coupled to a generator for producing electricity. In an effort to reduce pollution levels, government agencies have introduced new regulations requiring gas turbine engines to reduce emitted levels of emissions, including carbon monoxide (CO) and oxides of nitrogen (NOx). A common type of combustion, employed to comply with these new emissions requirements, is premix combustion, where fuel and compressed air are mixed together prior to ignition to form as homogeneous a mixture as possible and burning this mixture to produce lower emissions. While premixing fuel and compressed air prior to combustion has its advantages in terms of emissions, it also has certain disadvantages such as combustion instabilities and more specifically combustion dynamics.
In order to achieve the lowest possible emissions through premixed combustion, without the use of a catalyst, it is necessary to provide a fuel-lean mixture to the combustor. Generally, in combustors using fuel-lean mixtures, as the fuel-air mixture becomes more fuel rich (i.e. higher fuel-air ratio), the flame and combustion process becomes more stable. (Of course, if the fuel-air mixture becomes too fuel rich, the flame and combustion process becomes more unstable, and can lead to rich fuel blowout.) Therefore, fuel-lean mixtures tend to be more unstable given the lesser fuel content for a given amount of air. As a result, when fuel-lean mixtures are burned they tend to produce greater pressure fluctuations due to the unstable flame. A factor contributing to the unstable flame is the fuel-air ratio or more specifically, the amount of air mixing with a known amount of fuel. The amount of air entering into a combustion chamber can vary depending on how the air is directed towards the combustion chamber inlet. If the airflow is not uniform and not relatively free from swirl, the amount of air entering the combustor will fluctuate, thereby altering the fuel-air ratio, and adversely affecting combustion stability and NOx emissions.
An example of a gas turbine combustor of the prior art that employs premix combustion is shown in cross section in
Therefore, it is desired to provide a combustion system for a gas turbine wherein the geometry of the combustor provides a means for significantly reducing the tangential velocity, or swirl, for air directed to a combustion inlet so as to reduce combustion stability problems and NOx, and to reduce the overall pressure drop required across the combustor. Reducing the combustor pressure drop, will in turn improve the gas turbine efficiency, increase power output, and lower fuel operating cost.
An apparatus and method of providing a gas turbine combustor having increased combustion stability, lower NOx, reduced pressure drop across a gas turbine combustor, and higher efficiency is provided. A gas turbine combustor comprising a flow sleeve, combustion liner, at least one fuel nozzle, and a plurality of vanes fixed to the flow sleeve radially between the flow sleeve and combustion liner is disclosed. The plurality of vanes serve to mechanically direct a flow of air entering the region between the flow sleeve and combustion liner in a substantially axial direction, such that components of tangential velocity are removed thereby providing a more uniform flow of air to the combustion chamber and reducing the amount of pressure lost due attempting to straighten the airflow by pressure drop alone.
It is an object of the present invention to provide a gas turbine combustor having a reduced pressure drop across the combustor thereby improving the thermal efficiency of the gas turbine.
It is another object of the present invention to provide a gas turbine combustor having improved combustion stability and reduced dynamics by providing a more uniform air flow to the combustion chamber.
It is another object of the present invention to provide a gas turbine combustor having reduced NOx as compared to the gas turbine combustors of the prior art.
In accordance with these and other objects, which will become apparent hereinafter, the instant invention will now be described with particular reference to the accompanying drawings.
The preferred embodiment of the present invention will now be described in detail with particular reference to
An additional feature of flow sleeve 22 is plurality of vanes 30 that are fixed to flow sleeve 22 proximate plurality of first holes 25. Plurality of vanes 30 extend radially inward towards combustion liner 26 into first passage 27. The quantity of plurality of vanes 30 preferably corresponds equally to the quantity of plurality of first holes 25 as shown in
In addition to the apparatus described above, a method for reducing the pressure drop across a gas turbine combustor is disclosed that incorporates the combustion apparatus of the present invention. A method for reducing pressure drop across a combustor comprises the steps of providing a gas turbine combustor 21 comprising a flow sleeve 22 having a first end 23, a second end 24, and a plurality of first holes 25 located proximate second end 24. Combustor 21 also comprises combustion liner 26 located radially within flow sleeve 22, thereby forming first passage 27 therebetween, and at least one fuel nozzle 28 for injecting a fuel to mix with air in the combustion liner. Furthermore, combustor 21 comprises a plurality of vanes 30 fixed to flow sleeve 22 proximate plurality of first holes 25 and extending radially inward into first passage 27 towards combustion liner 26. Next, a flow of compressed air is directed through plurality of first holes 25, into first passage 27, and between plurality of vanes 30. The airflow is then straightened by the plurality of vanes 30 to significantly remove the tangential velocity component from the flow of compressed air and then directed in a substantially axial direction towards flow sleeve first end 23 in a more uniform pattern. As a result of the plurality of first holes 25 and plurality of vanes 30 mechanically straightening the passing airflow, pressure drop across combustor 21 from flow sleeve second end 24 to flow sleeve first end 23 is reduced. A lower pressure drop across flow sleeve 22 and first passage 27 results in more power output, higher thermal efficiency, and reduced fuel operating costs for the gas turbine in which the combustor of the present invention is used.
While the invention has been described in what is known as presently the preferred embodiment, it is to be understood that the invention is not to be limited to the disclosed embodiment but, on the contrary, is intended to cover various modifications and equivalent arrangements within the scope of the following claims.
