Patent Application
20130192243
The field of the invention relates generally to rotating machines and, more particularly, to turbine engine fuel nozzle assemblies.
At least some known turbine engines ignite a fuel-air mixture in a combustor to generate combustion gases that are channeled towards a turbine via a hot gas path. Known combustor assemblies include fuel nozzles that channel fuel to a combustion region of the combustor. The turbine converts thermal energy of the combustion gas stream to mechanical energy used to rotate a turbine shaft. Output of the turbine may be used to power a machine, for example, an electric generator, a compressor, or a pump.
Such known fuel nozzles are configured to inject and distribute a particular fuel blend to produce a predetermined flow rate through existing injection openings and jets defined within the fuel nozzles. Such known fuel nozzles are statically configured and include no features to respond to dynamic conditions, e.g., varying fuel heat content, fuel viscosities, and volumetric flow rates. If an owner/operator desires to switch fuels, the owner/operator will remove the turbine engine from service and replace the nozzles with alternative nozzles to accommodate long-term fuel switching. In some gas turbine fuel systems, fuel flow adjustment devices, e.g., adjustable fuel flow throttle valves, are positioned upstream of the fuel nozzles. In some other gas turbine fuel systems, multiple fuel supply circuits are installed. Both solutions facilitate accommodating fuel switching on the fly. However, these two solutions also increase the costs of installation and operation of the associated fuel systems.
In one aspect, a fuel nozzle assembly for use with a turbine engine is provided. The fuel nozzle assembly includes at least one fuel conduit coupled to at least one fuel source. The fuel nozzle assembly also includes at least one swirler that includes at least one wall having a porous portion. The at least one wall is coupled to the at least one fuel conduit. The porous portion is formed from a material having a porosity that facilitates fuel flow therethrough. At least one fuel flow path is thereby defined through the porous portion of the at least one wall.
In a further aspect, a method of operating a turbine engine is provided. The turbine engine includes at least one fuel conduit coupled to at least one fuel nozzle that includes at least one swirler having at least one porous portion. The porous portion covers at least a portion of the at least one fuel conduit opening. The porous portion is formed from a material having a porosity that facilitates fuel flow therethrough. The method includes channeling a fuel from at least one fuel source to the at least one fuel conduit. The method also includes channeling the fuel through the at least one porous portion into a combustor.
In another aspect, a gas turbine engine is provided. The gas turbine engine includes at least one combustor and at least one fuel nozzle assembly coupled to the at least one combustor. The at least one fuel nozzle assembly includes at least one fuel conduit coupled to at least one fuel source. The at least one fuel nozzle assembly also includes at least one swirler including at least one wall having a porous portion. The porous portion is formed from a material having a porosity that facilitates fuel flow therethrough. The at least one wall is coupled to the at least one fuel conduit, thereby defining at least one fuel flow path through the porous portion of the at least one wall.
These and other features, aspects, and advantages of the present invention will become better understood when the following detailed description is read with reference to the accompanying drawings in which like characters represent like parts throughout the drawings, wherein:
Unless otherwise indicated, the drawings provided herein are meant to illustrate key inventive features of the invention. These key inventive features are believed to be applicable in a wide variety of systems comprising one or more embodiments of the invention. As such, the drawings are not meant to include all conventional features known by those of ordinary skill in the art to be required for the practice of the invention.
In the following specification and the claims, reference will be made to a number of terms, which shall be defined to have the following meanings.
The singular forms “a”, “an”, and “the” include plural references unless the context clearly dictates otherwise.
“Optional” or “optionally” means that the subsequently described event or circumstance may or may not occur, and that the description includes instances where the event occurs and instances where it does not.
Approximating language, as used herein throughout the specification and claims, may be applied to modify any quantitative representation that could permissibly vary without resulting in a change in the basic function to which it is related. Accordingly, a value modified by a term or terms, such as “about” and “substantially”, are not to be limited to the precise value specified. In at least some instances, the approximating language may correspond to the precision of an instrument for measuring the value. Here and throughout the specification and claims, range limitations may be combined and/or interchanged, such ranges are identified and include all the sub-ranges contained therein unless context or language indicates otherwise.
In the exemplary embodiment, combustor section 106 includes a plurality of combustor assemblies, i.e., combustors 116 that are each coupled in flow communication with compressor section 104. Combustor section 106 also includes at least one fuel nozzle assembly 118. Each combustor 116 is in flow communication with at least one fuel nozzle assembly 118. Moreover, in the exemplary embodiment, turbine section 108 and compressor section 104 are rotatably coupled to a load 120 via drive shaft 114. For example, load 120 may include, without limitation, an electrical generator and/or a mechanical drive application, e.g., a pump. Alternatively, gas turbine engine 100 may be an aircraft engine. In the exemplary embodiment, compressor section 104 includes at least one compressor blade assembly 122. Also, in the exemplary embodiment, turbine section 108 includes at least one turbine blade, i.e., bucket 124. Each compressor blade assembly 122 and each turbine bucket 124 is coupled to rotor assembly 112, or, more specifically, compressor drive shaft 114 and turbine drive shaft 115.
In operation, air intake section 102 channels air 150 towards compressor section 104. Compressor section 104 compresses inlet air 150 to higher pressures and temperatures prior to discharging compressed air 152 towards combustor section 106. Compressed air 152 is channeled to fuel nozzle assembly 118, mixed with fuel (not shown), and burned within each combustor 116 to generate combustion gases 154 that are channeled downstream towards turbine section 108. Combustion gases 154 generated within combustors 116 are channeled downstream towards turbine section 108. After impinging turbine bucket 124, thermal energy is converted to mechanical rotational energy that is used to drive rotor assembly 112. Turbine section 108 drives compressor section 104 and/or load 120 via drive shafts 114 and 115, and exhaust gases 156 are discharged through exhaust section 110 to ambient atmosphere.
Also, in the exemplary embodiment, fuel nozzle assembly 118 includes a swirler 222. Swirler 222 includes a plurality of vanes 224 disposed within main air conduit 220. Swirler vanes 224 are coupled in flow communication with main premix fuel conduit 216.
Further, in the exemplary embodiment, diffusion fuel port 210, diffusion fuel conduit 212, main premix fuel port 214, and main premix fuel conduit 216 are coupled in flow communication with a plurality of gaseous fuel sources (none shown) to enable one or more gaseous fuels to be selectively channeled to fuel nozzle assembly. In the exemplary embodiment, the fuels are a carbonaceous gas such as, but not limited to, a natural gas and a syngas. Alternatively, the fuels supplied to fuel nozzle assembly 118 may be any gaseous fuels that enable operation of fuel nozzle assembly 118 and gas turbine engine 100 as described herein.
Moreover, in the exemplary embodiment, fuel nozzle assembly 118 is coupled in flow communication with a cooling/purge air source (not shown) via cooling/purge air connection 208. The cooling/purge air source may include a portion of pressurized air 152 channeled from compressor section 104 (shown in
Also, in the exemplary embodiment, at least one portion of vane walls 230, i.e., at least one of portions 234 and 236 of vane walls 230, are formed from a porous material. In the exemplary embodiment, both portions 234 and 236 are porous. Specifically, portions 234 and 236 of vane walls 230 are formed from a material having a predetermined porosity that facilitates fuel flow therethrough, i.e., facilitates fuel effusion into main air conduit 220. Alternatively, only one of portions 234 and 236 are porous. Also, alternatively, other predetermined portions of vane walls 230 are formed from a porous material, either with or without porous portions 234 and 236.
Further, in the exemplary embodiment, each of porous portions 234 and 236 have a predetermined porosity value that is substantially constant radially and axially. Alternatively, each of porous portions 234 and 236 may have dissimilar predetermined porosity values that vary radially, axially, and with respect to each other.
Moreover, in the exemplary embodiment, porous portions 234 and 236 are directly opposite each other in an orientation that is parallel to axial centerline 202 (shown in
Porous material 300 is formed to define a predetermined porosity, or void fraction, as a measure of the void spaces in material 300, i.e., a fraction of the volume of voids over the total volume, between 0 and 1, or as a percentage, between 0 and 100%. In the exemplary embodiment, the void fraction is within a range of approximately 20% to approximately 70%. Alternatively, the void fraction is any value that enables operation of fuel nozzle assembly 118 and gas turbine engine 100 (shown in
Referring to
Referring to
In the exemplary embodiment, the predetermined porosity values of opposing porous portions 234 and 236 facilitate decreasing the formation of concentrated jets of fuel being injected into main air conduit 220 with fuels having different BTU values. Furthermore, reliance of fuel nozzle assemblies 118 (shown in
For at least some fuels, the characteristics of fuel injection into gas turbine engine 100 is more consistent to reduce effects induced by higher and lower differential pressures, varying viscosities, and pressure drops in the fuel flow paths. Also, the natural flow-restrictive features of porous material 300 facilitate a predetermined distribution of the fuel into main air conduit 220 within predetermined tolerances, including, without limitation, a substantially even distribution. Control of flow rates by use of fuel control devices upstream of, and external to, gas turbine engine 100 facilitates maintaining the fuel backpressure within predetermined ranges for each fuel. Moreover, in addition to control of fuel flow rates, effects from undesirable conditions such as fuel flow imbalances between multiple fuel nozzles 118 (shown in
Method 500 also includes transitioning 506 from a first fuel to a second fuel. The first fuel is channeled 508 from a first fuel source, wherein the first fuel has a first set of characteristics. The second fuel is channeled 510 from a second fuel source, wherein the second fuel has a second set of characteristics. Fuel flow from the first fuel source is decreased 512. Fuel flow from the second fuel source is increased 514. A power output of gas turbine engine 100 is held 516 substantially constant by varying 518 a fuel injection rate at least partially as a function of the first set of fuel characteristics, the second set of fuel characteristics, and a porosity of porous portions 234 and 236.
The percentages of the first and second fuels channeled into combustor 116 through porous portions 234 and 236 of swirler vanes 224 varies within a range extending between 0% and 100%. Also, fuel blends burned within combustor 116 are not limited to two different fuels.
The above-described fuel nozzle assembly provides a cost-effective method for increasing reliability and decreasing disruptions of operation of gas turbine engines. Specifically, the devices, systems, and methods described herein provide swirler vane walls of fuel nozzle assemblies having predetermined porosity values that facilitate switching fuels and modulating components of fuel blends on the fly. The devices, systems, and methods described herein decrease fuel injection variation and challenges with discrete concentrated jets when utilizing fuels having different BTU values. Furthermore, the devices, systems, and methods described herein decrease a reliance of the fuel nozzle assemblies on jet-in-crossflow mixing prior to channeling the fuel into the associated combustors. Moreover, the devices, systems, and methods described herein facilitate the variability of fuel flow rates for different fuels, regardless of different fuel blends requiring different flow rates, due to the predetermined porosity values of the swirler vane walls and the diffusive nature of the porous material.
Furthermore, the devices, systems, and methods described herein facilitate effective combustion of a wide range of fuels without maintenance outages to modify the fuel nozzle assemblies and the upstream fuel conduits. The devices, systems, and methods described herein facilitate, for at least some fuels, controlling the effects of the varying characteristics of fuel injection into a gas turbine engine, i.e., making the characteristics of fuel injection more consistent to reduce effects induced by higher and lower differential pressures, varying viscosities, and pressure drops in the fuel flow paths. Also, the devices, systems, and methods described herein facilitate using the natural flow-restrictive features of the porous material to provide for a predetermined distribution of fuel. Furthermore, the devices, systems, and methods described herein facilitate controlling the fuel flow rates by use of fuel control devices upstream of, and external to, a gas turbine engine to facilitate maintaining the fuel backpressure within predetermined ranges for each fuel. Moreover, in addition to control of fuel flow rates, effects from undesirable conditions such as fuel flow imbalances, burning fuels with out-of-specification characteristics, and out-of-tolerance distribution profiles are reduced.
An exemplary technical effect of the methods, systems, and apparatus described herein includes at least one of: (a) increasing reliability and decreasing disruptions of operation of gas turbine engines while switching fuels on the fly and modulating components of fuel blends; (b) providing swirler vane walls of fuel nozzle assemblies having predetermined porosity values that facilitate decreasing fuel injection variation and challenges with discrete concentrated jets when utilizing fuels having different BTU values; (c) providing swirler vane walls of fuel nozzle assemblies having predetermined porosity values that facilitate decreasing a reliance of the fuel nozzle assemblies on jet-in-crossflow mixing prior to channeling the fuel into the associated combustors; (d) providing swirler vane walls of fuel nozzle assemblies having predetermined porosity values that facilitate operation of gas turbine engines for different fuels requiring different flow rates; (e) providing swirler vane walls of fuel nozzle assemblies having predetermined porosity values that facilitate effective combustion of a wide range of fuels and fuel blends without maintenance outages to modify the fuel nozzle assemblies and the upstream conduits; (f) providing swirler vane walls of fuel nozzle assemblies having predetermined porosity values that facilitate controlling the effects of the varying characteristics of fuel injection into a gas turbine engine, i.e., making the characteristics of fuel injection more consistent to reduce effects induced by higher and lower differential pressures, varying viscosities, and pressure drops in the fuel flow paths; (g) providing swirler vane walls of fuel nozzle assemblies having predetermined porosity values that facilitate using the natural flow-restrictive features of the porous material to control the fuel flow rates by use of fuel control devices upstream of and external to a gas turbine engine to maintain the fuel backpressure within predetermined ranges for each fuel; and (h) providing swirler vane walls of fuel nozzle assemblies having predetermined porosity values that facilitate reducing effects of fuel flow imbalances, out-of-specification fuel flow, and out-of-tolerance distribution profiles.
Exemplary embodiments of fuel nozzle assemblies for gas turbine engines and methods for operating are described above in detail. The fuel nozzle assemblies and methods of operating such assemblies are not limited to the specific embodiments described herein, but rather, components of systems and/or steps of the methods may be utilized independently and separately from other components and/or steps described herein. For example, the methods may also be used in combination with other combustion systems and methods, and are not limited to practice with only the gas turbine systems and fuel nozzle assemblies and methods as described herein. Rather, the exemplary embodiment can be implemented and utilized in connection with many other combustion applications.
Although specific features of various embodiments of the invention may be shown in some drawings and not in others, this is for convenience only. In accordance with the principles of the invention, any feature of a drawing may be referenced and/or claimed in combination with any feature of any other drawing.
This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. 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 language of the claims.
This invention was made with Government support under contract number DE-FC26-08NT05868 awarded by the Department of Energy (DOE). The Government may have certain rights in this invention.
