GB2094464A - Gas turbine combustor - Google Patents

Abstract

A combustor (10) is capable of reducing the noxious emissions such as fuel bound and thermal nitrogen oxide products during combustion of high nitrogen bearing and high aromatic content fuels. It includes a plurality of substantially concentric pipes (11 to 16) defining annular passages (12a to 16a) with annular venturi shaped divergent nozzles (20 to 24) to facilitate fast mixing of axially supplied air between adjacent annular passages. The longitudinal spacing between at least two adjacent nozzles defines first and second divergent cavities (30, 40). A fuel rich toroidal vortex (T) is formed in proximity to a central fuel jet in the first cavity (30) and advantageously converts fuel bound nitrogen to N2. A fuel lean toroidal vortex (T') formed in the second cavity (40) mixes hot combustion products with additional gaseous reactant to complete the combustion while avoiding locally high temperatures, and thus thermal NOx formation. A ring of jet nozzles (42) radially injects relatively small amounts of high pressure gaseous reactant or steam to form a throat to separate and stabilize the vortices. Alternatively, the pipes extending between the two cavities can include a convergent and divergent portion forming the throat. <IMAGE>

Description

SPECIFICATION Gas turbine combustor This invention relates generally to gas turbine combustors and, more particularly, to two-stage combustors capable of developing separate fuel rich and fuel lean zones for improved combustion and to minimize formation of nitrogen oxide (NOx) products.

Combustors are used in gas turbines for developing high pressure gases used in the generation of turbine power. In such turbine systems, gaseous reactant and fuel supplied by a compressor to a combustion chamber of the combustor are ignited and discharged into the inlet side of a turbine. The present practice is to use relatively refined fuels, such as kerosene or diesel fuels, or natural gas, that previously were relatively easily available; the gaseous reactant may be air, oxygen or oxygen enriched air, or carbon dioxide.By mixing and igniting the fuel and gaseous reactant, high volumetric heat release rates can be obtained under turbulent conditions by matching the concentrations and directions of fuel and gaseous reactant flow in a manner enabling high fuel concentration regions to overlap with regions of large shear stresses in the gaseous reactant flow, as disclosed in British Patent No. 1,099,959, issued January 17, 1968.

It is recognized as desirable, especially in light of the energy shortage, to be able to use lower grade fuels, such as high nitrogen bearing, high aromatic content petroleum fuels, shale oils and coal liquids, for turbine power.

The major problems in addition to efficiency and proper mixing of the gases and these fuels, are fiame stabilization, elimination of pulsation and noise, and control of pollutant emissions, especially carbonaceous particulates and nitrogen oxides (NOx). Nitrogen oxides emitted from combustion processes have two main sources; namely the fixation of atmospheric nitrogen from the combustion air at high temperatures, and the conversion of organically bound nitrogen compounds in the fuel to NOx. When the nitrogen content of the fuel exceeds 0.1% by weight, the fuel bound nitrogen plays an increasingly significant role in the emission of NOx. However, the laws governing formation of NOx from these two major sources are quite different.For example, the formation of NOx from atmospheric nitrogen is primarily dependent upon combustion temperature, and generally referred to as "thermal NOx"; whereas the rate of formation of NOx from organically bound nitrogen in the fuel, generally referred to as "fuel NOx," is largely dependent upon local fuel-air mixture ratios and to a lesser extent upon temperature.

To minimize conversion of fuel bound nitrogen to NOx, it is necessary to first pyrolyse the fuel by heating it in an oxygen deficient environment, followed by admixing the combustion products and combustion air to complete the combustion process. Recent research has shown that given fuel rich conditions and sufficient residence time and temperature in the first or pyrolysis stage of the combustion process, fuel bound nitrogen may be rendered innocuous for NOx formation in the fuel lean second stage. This occurs through conversion to molecular nitrogen (N2) in the fuel rich first stage. However, care has to be taken when the rest of the combustion air is admixed to avoid locally high temperatures resulting in the formation of thermal NOx.This is achieved by admixing of combustion air and products of pyrolysis such that the temperature of the mixture is initially reduced by rapid mixing. This effects quenching of the reactions that would otherwise lead to the formation of thermal NOx.

Downstream, a temperature rise occurs due to the up take of the oxygen by the pyrolysis products and exothermic combustion reactions. To effectuate these conditions, the temperature history of the mixture has to be closely controlled to insure that the combustion of soot and hydrocarbons may proceed to completion within the residence time in the combustor while maintaining temperatures in the lean stage below 1 6000K.

It is accordingly an object of the present invention to provide a gas turbine combustor capable of minimizing the formation of nitrogen oxide products by tailoring the mixing and temperature history of the fuel according to known thermodynamic and chemical kinetic requirements of the combustion process.

The invention resides in a gas turbine combustor for reducing the emission of fuel bound and thermal nitrogen oxide (NOx) products during combustion of high nitrogen bearing and high aromatic content fuels, characterized in that the combustor comprises a plurality of substantially concentric pipes defining annular passages having central and annular openings located at an end of said pipes for receiving fuel and swirling gaseous reactant; a plurality of substantially concentric annular divergent nozzles positioned within the passages, the longitudinal spacing between at least two nozzles being such as to define first and second divergent cavities communicating with each other, said second cavity being positioned downstream from the first cavity, said spacing between said nozzles forming a fuel rich and fuel lean toroidal vortex respectively in said first and second cavities; and throat means positioned between the first and second cavities for separating the fuel rich and fuel lean vortices, said throat means being operable to control fast admixing of gaseous reactant supplied to the second cavity with combustion products introduced from the first cavity, to minimize formation of fuel bound and thermal NOx products during combustion.

The gas turbine combustor according to the present invention comprises first and second combustion zones wherein a first fuel rich zone minimizes the conversion of fuel bound nitrogen to NOx and the second fuel lean zone fast mixes the combustion products from the first zone with combustion air at temperatures sufficiently low to prevent formation of thermal NOx. In the combustor, cooling of the combustor walls is recuperative so that heat loss in the fuel rich zone is reduced, without using any part of the gaseous reactant for film cooling. Good control of the flow and mixing pattern is achieved while minimizing the pressure drop through the combuston In addition, the combustor maintains temperatures sufficiently high for complete combustion without the formation of NOx products.

In accordance with a preferred embodiment of the present invention, the gas turbine combustor comprises a plurality of substantially concentric pipes defining annular pasages having central and annular openings located at one end of the pipes for receiving fuel and swirling gaseous reactant. A plura(ity of substantially concentric annular divergent nozzles are positioned within the passages, and the longitudinal spacing between at leasttwo adjacent nozzles defines first and second divergent cavities formed by the nozzle ends. The first cavity is formed in proximity to a central fuel injector to create a first stage fuel rich zone, and the second divergent cavity is positioned downstream from the first cavity forming a second stage fuel lean zone, whereby complete combustion is effected.The spacing between the nozzles within the first and second cavities is conducive to forming fuel rich and fuel lean toroidal vortices respectively in each cavity.

Preferably,the axial spacing of adjacent nozzles forming the first cavity increases relative to the radial distance from the combustor axis. This geometrical pattern forms an envelope with substantially concave boundaries; whereas, constant axial spacing of nozzles forming the second cavity defines substantially straight line boundaries.

Throat means is positioned between the first and second cavities of the combustorfor separating and reinforcing the fuel rich and fuel lean vortices. In the preferred embodiment, such means includes a ring jet circumferentially positioned around the combustor for radially injecting small amounts of high pressure gaseous reactant directly into the fuel rich vortex in proximity to a stagnation point of the vortex. In embodiment of the present invention, such means preferably includes a throat section of the concentric pipe located between the two adjacent nozzles, The throat section includes a convergent portion, and a divergent portion integrally formed therewith and downstream from the convergent portion. This structure provides separating and reinforcing action to the formation of the gaseous toroidal vortices.

Swirl generating means is positioned in the concentric passages for imparting a swirl velocity component to gaseous reactants axially supplied through the annular passages, enabling rotation of the gaseous reactant for forming the toroidal vortices. Such means preferably includes a plurality of turbine stator-type guide vanes fixedly attached at spaced circumferential intervals within the annular passages at a predetermined vane angle.

The guide vane angles can be adjusted for achieving greatest swirl velocity in an innermost annular passage communicating with the first cavity, and a swirl velocity gradually decreasing with increasing radial distance from the longitudinal axis of the combustor.

The divergent nozzles of the combustor are preferably formed buy a ring having a venturi shaped axial section. This geometry facilitates fast mixing of axially supplied air between adjacent annular passages for maximum combustion efficiency, and thus minimum pollution.

The invention will become radially apparent from the following description of exemplary embodiments thereof when taken in conjunction with the accompanying drawings in which: Figure 1 is a schematic view of the gas turbine combustor according to a preferred embodiment of the present invention showing the formation of fuel rich and fuel lean toroidal votices respectively in the first and second combustion cavities; Figure 2 is a schematic view of a second embodiment according to the present invention showing the use of a convergent-divergent throat section for separating and strengthening the toroidal vortices in the first and second combustion cavities; and Figure 3 is an enlarged side view partially broken away showing in addition detail the positioning of swirl vanes in the annular passages between the concentric pipes.

Referring first to Figure 1, a combustor 10 is shown comprising six pipes 11-16 of progressively larger diameter. These pipes may be mounted in a conventional manner (not shown) in a heat generating system, a power turbine or similar systems. The overlapping, substantially concentric alignment of the pipes 11-16 defines a central passage 11 a and annular passages 1 2a-1 6a extending longitudinally between the corresponding pipe walls. Each of central and annular passages 11 a-1 6a respectively defines a central intake opening and annular intake openings formed at one end of the pipes (note flow arrows in Figure 1).Annular divergent nozzles 20-24 are respectively positioned within the outlet openings of the pipes 11-16 along the inner end of the inner pipes. These nozzles serve to form gaseous envelopes including toroidal vortices (see Figure 1), thus defining first and second combustion cavities or stages 30, 40, respectively. Fuel jet or inlet nozzle 31 is positioned within the central opening along combustor longitudinal axis L for supplying fuel to first cavity 30.

First cavity 30 forms a fuel rich stage of combustor 10 extending forwardly from fuel jet 31 along divergent nozzles 20-22. As shown, the axial spacing of these nozzles increases in relation to their radial distance from combustor axis L to define a divergent cavity with a substantially concave outer boundary. Second cavity 40 forms a fuel lean stage of combustor 10 along divergent nozzles 22-24. These secondary divergent nozzles are equally spaced apart in relation to their radial distance from burner axis L to define a second divergent cavity having a substantially straight-line outer boundary. This second stage is immediately downstream from first cavity 30.As shown in Figure 1, each of first and second cavities 30, 40 includes three divergent nozzles and wherein outermost divergent nozzle 22 of the first cavity substantially defines the innermost nozzle of the second cavity.

As shown in Figure 3, a plurality of turbine stator-type guide vanes 45 are positioned at spaced circumferential intervals in each of the annular openings for imparting a swirl velocity component to gaseous reactant entering the passages 1 2a-1 4a. The intake reactant may be supplied by a compressor (not shown). The rotation of the gaseous reactant about combustor axis L is a beneficial factor in the increased ,efficiency of combustion and the control of the gaseous temperatures in the two stages to reduce pollution in the exhaust, as will be explained in further detail below. Guide vanes 45 are secured to the inner pipe walls of each pair of pipes defining one of annular passages 1 2a-1 6a.

Guide vanes 45 preferably have a fixed blade angle A (see Figure 3) for rotating gaseous reactants about burner axis L. A more complete discussion of guide vanes 45 may be found in Combustion Aerodynamics by J. M. Beer and N. A.

Chigier, Eisevier, 1972, Chapter 5.

In operation, liquid, gaseous, or slurry fuel is injected into the first cavity 30 through fuel jet or nozzle 31 and mixes with gaseous reactant supplied through divergent nozzles 20-22 of the first stage. The highly swirling gaseous reactant flow in combination with the divergence within first cavity 30 is operable to generate the toroidal vortex pattern, as indicated by streamlines T, (Figure 1). In the second cavity 40, a second toroidal vortex with streamlines T' is generated within the envelope of the reactant entering the cavity through the annular passages 1 5a, 1 6a.

Each toroidal vortex extends longitudinally within a cavity and has a recirculating flow pattern along combustor axis L in the direction of fuel jet 31. A stragnation pressure area P exists slightly downstream of each toroidal vortex T, T'. To achieve proper flame stabilization and combustion in first cavity 30, the axial spacing between divergent nozzles 21, 22 must be sufficiently large for maintaining proper separation of the vortices T, T' in each cavity 30, 40, as discussed below. The first of these vortices constitutes the fuel rich stage of combustor 10 consisting of the fuel introduced along burner axis L and a proportion of the stoichiometric combustion air.Typically twothirds of the stoichiometric combustion air is introduced through the three innermost pipes 1 2-14. The vigorous stirring in this zone is essential for the fast vaporization of the liquid fuel, the efficient conversion of fuel bound nitrogen to N2, and also to avoid excessive formation of soot in the fuel rich zone. The second tornidal vortex T' formed in second cavity 40 embodies a fuel lean combustion stage in which the combustion products of the first stage are rapidly cooled to quench the thermal NOx formation reaction while maintaining the mixture temperature high enough for completing the combustion of carbon monoxide, hydrocarbons and soot leaving first cavity 30.

The cooling of the pipe walls (i.e. the sections of pipes 12-1 6 between divergent nozzles 20--24) is recuperative, enabling the total amount of gaseous reactant to cool the walls by flowing past them and return heat into the combustion system of first and second cavities 30, 40. This envelope surrounding the vortex reduces heat loss from the fuel rich stage which is desirable since high temperatures assist in speeding up the chemical reactions converting the fuel bound nitrogen to N2. All of the gaseous reactant enters axially effectively cooling the pipe walls.There is no need to use part of the gaseous reactant as "film cooling" for the walls, thus enabling the total amount of gaseous reactant to be available for the efficient management of the flow and mixing pattern in combustor 1 0. The feature of providing good control over the flow and mixing pattern with a simple burner geometry further enables the pressure drop across the combustor to be maintained at lower levels than in conventional combustors operating at corresponding performance levels.

The gaseous reactant necessary for completing combustion and reducing temperature in the fuel lean zone of second cavity 40 is provided through divergent nozzles 22-24. Fast mixing between this gaseous reactant and the products of the fuel rich zone result in lowering the mixing temperature to below 16000K, necessary for ensuring that little or no thermal NOx is formed, yet operable to maintain the temperature sufficiently high to burn the combustibles. High turbulent shear stresses arising between adjacent divergent nozzles result in uniform distribution of fluid properties, such as gas temperature, across the cross section of combustor 1 0, which is advantageous for gas turbine applications.If necessary, additional fuel, whether liquid, gaseous, or a slurry may be introduced at other positions along the burner, either axially through a ring jet (not shown) in the pipes, or tangentially through one or more of the pipe walls between adjacent divergent nozzles.

Forthe purpose of stabilizing the toroidal vortices the further strengthening the recirculating flow of the fuel rich vortex, throat means is provided for increasing stagnation pressure in area P. As shown in Figure 1, such means preferably includes a ring 42 of jets extending around pipe 1 4 between divergent nozzles 21, 22. Pressurized air is injected radially inward through ring jets 42 in this stagnation region P within the fuel rich toroidal vortex T. After combustion in the fuel rich vortex T, the combustion products from first cavity 30 pass downstream into second cavity 40 to complete combustion in the fuel lean vortex 40.

Figure 2 shows a second embodiment of the present invention, wherein an additional pipe 14' is mounted between pipes 14 and 1 5. A throat section of pipe 14' is located between longitudinally spaced, adjacent nozzles 21, 22 (defining first and second cavities 30, 40). The throat is provided with annular convergent wall sections 1 4a' and divergent wall sections 1 4b', thus defining a throat passage capable of separating the fuel rich and fuel lean vortices by increasing stagnation pressure at area P and reinforcing the recirculating flow of the fuel rich vortex. The feature of forming the throat in this manner also improves fast admixing of air in second cavity 40 with combustion products from first cavity 30 to quench thermal NOx formation reactions in the second cavity.In addition, strengthening of the recirculatina luel rich vortex is operable to return hot combustion products for mixing with fresh fuel to ensure flame stability.

To facilitate fast mixing between gaseous reactant supplied through adjacent annular passages 1 2a-1 6a, divergent nozzles 20-24 are contoured with venturi shaped axial sections.

As shown in Figure 1 and 2, each divergent nozzle 20-24 is formed of a ring having a section converging inwardly a short distance to a minimum inside diameter and then diverging gradually toward the exhaust openings. The adjacent pipe ends of the annular exhaust openings are preferably flared to continue the divergence of each nozzle.

To increase the strength of the fuel rich vortex recirculation flow it is desirable to adjust the angle of guide vanes 45 to achieve a highest swirl velocity in the innermost annular passage. The swirl velocity then decreases gradually with increasing radial distance from burner axis L.

To improve the recirculation flow of the fuel rich toroidal vortex in first cavity 30, the axial distance between adjacent nozzles increases radially from burner axis L to define a concavely shaped envelope within the divergent cavity. The curved shape extends along the tips of divergent nozzles as shown by projection line C.

By axially spacing divergent nozzles 22-24 in second cavity 40 to achieve a frusto-conical contour extending along the nozzles (shown by straight projection line C'), greater control over thermal NOx formation is achieved.

The foregoing description of a preferred embodiment of the invention has been presented for purposes of illustration and description. It is not intended to be exhaustive or to limit the invention to the precise form disclosed, and obviously many modifications and variations are possible in light of the above teaching. This embodiment was chosen and described in order to best explain the principles of the invention and as practicable application to thereby enable others skilled in the art to best utilize the invention in various embodiments and with various modifications as are suited to the particular use contemplated. It is intended that the scope of the invention be defined by the claims appended hereto.

Claims (9)

1. A gas turbine combustor for reducing the emission of fuel bound and thermal nitrogen oxide (NOx) products during combustion of high nitrogen bearing and high aromatic content fuels, characterized in that the combustor comprises a plurality of substantially concentric pipes defining annular passages having central and annular openings located at an end of said pipes for receiving fuel and swirling gaseous reactant; a plurality of substantially concentric annular divergent nozzles positioned within the passages, the longitudinal spacing between at least two nozzles being such as to define first and second divergent cavities communicating with each other, said second cavity being positioned downstream from the first cavity, said spacing between said nozzles forming a fuel rich and fuel lean toroidal vortex respectively in said first and second cavities; and throat means positioned between the first and second cavities for separating the fuel rich and fuel lean vortices, said throat means being operable to control fast admixing of gaseous reactant supplied to the second cavity with combustion products introduced from the first cavity, to minimize formation of fuel bound and thermal NOx products during combustion.

2. A gas turbine combustor as defined in claim 1, characterized in that it further includes swirl generating means positioned in the annular passages for imparting a swirl velocity component to gaseous reactant, thereby enabling rotation of the gaseous reactant about a longitudinal combustor axis.

3. A gas turbine combustor as defined in claim 2, characterized in that said swirl generating means includes a plurality of guide vanes fixedly attached at spaced circumferential intervals within the annular passages at a predetermined vane angle.

4. A gas turbine combustor as defined in claim 3, characterized in that the predetermined angle of the guide vanes is selected for achieving greatest swirl velocity in an innermost annular passage communicating with the first cavity, and a swirl velocity gradually decreasing with increasing radial distance from the longitudinal axis of the combustor, thereby enabling formation of the toroidal vortices to minimize fuel bound NOx products.

5. A gas turbine combustor as defined in claim 1, 2, 3 or 4, characterized in that each of said divergent nozzles is formed by a ring having a venturi shaped axial section to facilitate fast mixing between adjacent annular passages.

6. A gas turbine combustor as defined in any one of the preceding claims, characterized in that, said throat means includes a pipe section of a concentric pipe located between the two adjacent nozzles longitudinally spaced to define said first and second cavities, said pipe section having a convergent section and a divergent section.

7. A gas turbine combustor as defined in claim 1, 2, 3, 4 or 5, characterized in that said throat means includes means for:radially injecting relatively small amounts of high pressure fluid downstream and in proximity to a stagnation point of the fuel rich vortex, thereby stabilizing said vortex by increasing the stagnation pressure and further strengthening the recirculating flow of the fuel rich vortex.

8. A gas turbine combustor-as defined in any one of the preceding claims, characterized in that some of said concentric annular divergent nozzles defining the second divergent cavity receive gaseous reactant for completing the combustion process by fast mixing with combustion products supplied from the fuel rich vortex through said throat means, said mixing occurring such that the mixing temperature is below approximately 1 600 degrees K, thereby minimizing the formation of thermal NOx.

9. A gas turbine combustor as defined in any of the preceding claims, characterized in that, said divergent nozzles forming the first cavity are axially spaced to define an envelope with concave boundaries, said divergent nozzles forming the second cavity being axially spaced to define an envelope with straight line boundaries.