GB2635680A

GB2635680A - Combustor for gas turbine engine

Info

Publication number: GB2635680A
Application number: GB2317718.1A
Authority: GB
Inventors: Beita Jadeed
Original assignee: Siemens Energy Global GmbH and Co KG
Current assignee: Siemens Energy Global GmbH and Co KG
Priority date: 2023-11-20
Filing date: 2023-11-20
Publication date: 2025-05-28
Also published as: WO2025108672A1

Abstract

A combustor for a gas turbine engine comprises a radial swirler 202 coupled to receive respective flows of reactants including a flow of air 204 and a flow of main fuel 206. A premixing chamber 208 defines a premixing annulus 210 fluidly coupled to receive a swirling mix of the respective flows of reactants from the radial swirler. A pilot burner 212 has a pilot burner body including a pilot burner edge 214 having a plurality of pilot fuel injection orifices 216 arranged to inject pilot fuel for combustion in a combustion chamber 224 disposed downstream from the premixing chamber. At least a portion of the pilot burner body defines a first wall 222 of the premixing annulus, and a portion of a combustor liner 225 defines a second wall 220 of the premixing annulus. The first and second walls are coaxial walls configured to gradually turn from a radial direction to an axial direction. The pilot burner body axially extends beyond a downstream edge of the radial swirler to a point where a radial expansion of the combustion chamber is initiated.

Description

COMBUSTOR FOR GAS TURBINE ENGINE

BACKGROUND

[0001] The present disclosure relates to the field of combustion, and, more specifically, to a combustor for a combustion engine, such as a gas turbine engine.

[0002] In the field of combustion involving turbomachinery, it is a typical goal to try inhibiting formation of emissions, such as emissions of nitrogen oxides (NOx) that, for example, may be caused by the high temperatures reached in a combustion chamber. Appropriate mixing of fuel and gas (e.g., air) is considered one factor for inhibiting such elevated temperatures and thus for reducing overall NOx emissions, for example.

[0003] It is desirable to have a gas turbine engine that can be fueled with a more reactive fuel (or fuel mixture), such as involving hydrogen fuel, hydrogen enriched fuel blends, liquefied petroleum gas (LPG) fuel, etc. However, it is challenging building a gas turbine engine that can flexibly operate with various fuels or fuel blends, and which is effective to appropriately inhibit NOx emissions and reduce risk of flashback notwithstanding of operation with a more reactive fuel or fuel blend.

BRIEF SUMMARY

[0004] In one aspect, a combustor includes a radial swirler coupled to receive respective flows of reactants including a flow of air and a flow of main fuel. A premixing chamber defines a premixing annulus fluidly coupled to receive a swirling mix of the respective flows of reactants from the radial swirler. A pilot burner has a pilot burner body including a pilot burner edge having a plurality of pilot fuel injection orifices arranged to inject pilot fuel for combustion in a combustion chamber disposed downstream from the premixing chamber. At least a portion of the pilot burner body defines a first wall of the premixing annulus, and a portion of a combustor liner defines a second wall of the premixing annulus. The first wall and the second wall of the premixing annulus are coaxial walls configured to gradually turn from a radial direction to an axial direction. The pilot burner body axially extends beyond a downstream edge of the radial swirler to a point where a radial expansion of the combustion chamber is initiated.

BRIEF DESCRIPTION OF THE DRAWINGS

[0005] To easily identify the discussion of any particular element or act, the most significant digit or digits in a reference number refer to the figure number in which that element is first introduced.

100061 FIG. I is a longitudinal cross-sectional view of one non-limiting example of a gas turbine engine taken along a plane that depicts a longitudinal axis or central axis.

[0007] FIG. 2 is a fragmentary, longitudinal cross-sectional view of one non-limiting example of a disclosed combustor.

DETAILED DESCRIPTION

[0008] During flashback, a flame can propagate in an upstream direction, and thus the flame can "flash back" onto combustor hardware. Sustained upstream propagation of flames can potentially result in substantial thermal damage to the combustor hardware and associated components. Hydrogen, and hydrogen enriched fuel blends, for example, tend to exhibit relatively faster kinetics and higher flame speeds compared to traditional turbine fuels, such as natural gas, thus making hydrogen, and hydrogen enriched fuel blends more susceptible to flashback.

[0009] At least in view of the foregoing considerations, disclosed embodiments feature a reliable and cost-effective approach for premixing of fuel and air by way of an improved premixing chamber featuring a swept design, as elaborated in greater detail below. This improved design provides at least two main technical benefits: (i) ensures imparting a relatively higher velocity to the reactants flowing within the premixing chamber to inhibit flashback; and (ii) inhibits flow separation along the walls of the premixing chamber and further inhibits the development of recirculation zones, which otherwise could result in boundary layer flashback. These benefits are believed to enhance operational fuel flexibility in connection with disclosed embodiments.

[0010] Bcforc any embodiments are explained in detail, it is to be understood that disclosed embodiments are not limited by the details of construction and the arrangement of components set forth in this description or illustrated in the following drawings. Disclosed embodiments are capable of being practiced or of being carried out in various ways. Also, it is to be understood that the phraseology and terminology used herein is for the purpose of description and should not be regarded as limiting.

[0011] Various technologies that pertain to systems and methods will now be described with reference to the drawings, where like reference numerals represent like elements throughout. The drawings discussed below, and the various embodiments used to describe the principles of the present disclosure in this patent document are by way of illustration only and should not be construed in any way to limit the scope of the disclosure. Those skilled in the art will understand that the principles of the present disclosure may be implemented in any suitably arranged apparatus. It is to be understood that functionality that is described as being carried out by certain system elements may be performed by multiple elements. Similarly, for instance, an element may be configured to perform functionality that is described as being carried out by multiple elements. The numerous innovative teachings of the present application will be described with reference to exemplary non-limiting embodiments.

[0012] It should be understood that the words or phrases used herein should be construed broadly, unless expressly limited in some examples. For example, the terms "including," "having," and "comprising," as well as derivatives thereof, mean inclusion without limitation. The singular forms "a", "an" and "the" are intended to include the plural forms as well, unless the context clearly indicates otherwise. Further, the term "and/or" as used herein refers to and encompasses any and all possible combinations of one or more of the associated listed items. The term "or" is inclusive, meaning and/or, unless the context clearly indicates otherwise. The phrases "associated with" and "associated therewith," as well as derivatives thereof, may mean to include, be included within, interconnect with, contain, be contained within, connect to or with, couple to or with, be communicable with, cooperate with, interleave, juxtapose, be proximate to, be bound to or with, have, have a property of, or the like. Furthermore, while multiple embodiments or constructions may be described herein, any features, methods, steps, components, etc. described with regard to one embodiment are equally applicable to other embodiments absent a specific statement to the contrary.

[0013] Also, terms such as "first", "second", "third" and so forth may be used herein to refer to various elements, information, functions, or acts, but should not he considered as limiting in any way. Rather these numeral adjectives are used to distinguish different elements, information, functions or acts from each other. For example, a first element, information, function, or act could be termed a second element, information, function, or act, and, similarly, a second element, information, function, or act could be termed a first element, information, function, or act, without departing from the scope of the present disclosure.

[0014] In addition, the term "adjacent to" may mean that an element is relatively near to but not in contact with a further element or that the element is in contact with the further portion unless the context clearly indicates otherwise. Further, the phrase "based on" is intended to mean "based, at least in part, on" unless explicitly stated otherwise. Terms "about" or "substantially" or like terms are intended to cover variations in a value that are within normal industry manufacturing tolerances for that dimension. If no industry standard is available, a variation of twenty percent would fall within the meaning of these terms unless otherwise stated.

[0015] FIG. 1 illustrates an example of a turbine engine 100 including a compressor section 104, a combustion section 102, and a turbine section 106 arranged along a central axis 122. The compressor section 104 includes a plurality of compressor stages 108 with each compressor stage 108 including a set of rotating blades 126 and a set of stationary vanes 124 or adjustable guide vanes. A rotor 128 supports the rotating blades 126 for rotation about the central axis 122 during operation. In some constructions, a single one-piece rotor 128 extends the length of the gas turbine engine 100 and is supported for rotation by a bearing at either end. In other constructions, the rotor 128 is assembled from several separate spools that are attached to one another or may include multiple disk sections that are attached via a bolt or plurality of bolts.

[0016] The compressor section 104 is in fluid communication with an inlet section 116 to allow the gas turbine engine 100 to draw atmospheric air into the compressor section 104. During operation of the gas turbine engine 100, the compressor section 104 draws in atmospheric air and compresses that air for delivery to the combustion section 102. The illustrated compressor section 104 is an example of one compressor section 104 with other arrangements and designs being possible.

[0017] In the illustrated construction, the combustion section 102 includes a plurality of separate combustors 112 that each operate to mix a flow of fuel with the compressed air from the compressor section 104 and to combust that air-fuel mixture to produce a flow of high temperature, high pressure combustion gases or exhaust gas 118. Of course, many other arrangements of the combustion section 102 are possible.

[0018] The turbine section 106 includes a plurality of turbine stages 110 with each turbine stage I 10 including a number of rotating blades and a number of stationary blades or vanes. The turbine stages 110 are arranged to receive the exhaust gas 118 from the combustion section 102 at a turbine inlet 114 and expand that gas to convert thermal and pressure energy into rotating or mechanical work. The turbine section 106 is connected to the compressor section 104 to drive the compressor section 104. For gas turbine engines 100 used for power generation or as prime movers, the turbine section 106 is also connected to a generator, pump, or other device to be driven. As with the compressor section 104, other designs and arrangements of the turbine section 106 are possible.

[0019] A control system 120 is coupled to the gas turbine engine 100 and operates to monitor various operating parameters and to control various operations of the gas turbine engine 100. In preferred constructions the control system 120 is typically micro-processor based and includes memory devices and data storage devices for collecting, analyzing, and storing data. In addition, the control system 120 provides output data to various devices including monitors, printers, indicators, and the like that allow users to interface with the control system 120 to provide inputs or adjustments. In the example of a power generation system, a user may input a power output set point and the control system 120 may adjust the various control inputs to achieve that power output in an efficient manner.

[0020] The control system 120 can control various operating parameters including, but not limited to variable inlet guide vane positions, fuel flow rates and pressures, engine speed, valve positions, generator load, and generator excitation. Of course, other applications may have fewer or more controllable devices. The control system 120 also monitors various parameters to assure that the gas turbine engine 100 is operating properly. Some parameters that are monitored may include inlet air temperature, compressor outlet temperature and pressure, combustor outlet temperature, fuel flow rate, generator power output, bearing temperature, and the like. Many of these measurements are displayed for the user and are logged for later review should such a review be necessary.

[0021] FIG. 2 is a fragmentary, longitudinal cross-sectional view of one non-limiting example of a disclosed combustor 200 that includes a radial swirler 202 coupled to receive respective flows of reactants comprising a flow of air 204 and a flow of main fuel 206 from a main burner (not shown). A premixing chamber 208 defines a premixing annulus 210 fluidly coupled to receive a swirling mix 207 of the respective flows of reactants from radial swirler 202.

[0022] A pilot burner 212 has a pilot burner body extending to a pilot burner edge 214 where a plurality of pilot fuel injection orifices 216 are arranged to inject pilot fuel 218 for combustion in a combustion chamber 224 disposed downstream from premixing annulus 210.

In one example embodiment, the plurality of pilot fuel injection orifices 216 is arranged to radially outwardly eject the pilot fuel from pilot burner 212.

100231 In one example embodiment, at least a portion of the pilot burner body defines a first wall 220 of the premixing annulus 210, and a portion of a combustor liner 225 defines a second wall 222 of the premixing annulus 210. In one example embodiment, the first wall 220 and the second wall 222 of the premixing annulus 210 are coaxial walls relative to one another configured to gradually turn from a radial direction to an axial direction. In one example embodiment, the plurality of pilot fuel injection orifices 216 is circumferentially disposed along a periphery of the first wall 220 of the premixing annulus 210 about pilot burner edge 214. In one example embodiment, the pilot burner body axially extends beyond a downstream face of radial swirler 202 to a point where a radial expansion of combustion chamber 224 is initiated.

[0024] In one example embodiment, a diameter of premixing annulus 210 is less than a diameter defined from side to side of combustor liner 225 at the point where the radial expansion of combustion chamber 224 is initiated.

[0025] In one example embodiment, premixing annulus 210 comprises a first cross sectional area at an inlet of premixing annulus 210, and further comprises a second cross sectional area at an outlet of premixing annulus 210, where a ratio of the first cross sectional area and the second cross sectional area is less than unity. The inlet of premixing annulus 210 is disposed at the location where premixing annulus 210 receives the output flows from radial swirler 202 and the outlet of premixing annulus 210 is disposed at the location where premixing annulus 210 conveys swirling mix 207 into combustion chamber 224. In one example embodiment, the ratio of the first cross sectional area and the second cross sectional area is in a range from 0.7 to 0.9.

100261 In one example embodiment, the first cross sectional area and the second cross sectional area is each respectively smaller relative to a size of a cross-sectional area (e.g., circular area) defined by combustor liner 225 at the location where the radial expansion of combustion chamber 224 takes place.

[0027] In one example embodiment, a respective shape of the first wall 220 of premixing annulus 210 and a respective shape of the second wall 222 of premixing annulus 210 each is defined by a parabolic function. In one example embodiment, a respective shape of the first wall 220 of premixing annulus 210 and a respective shape of the second wall 222 of premixing annulus 210 each is defined by a polynomial function. In one example embodiment, a respective shape of the first wall 220 of premixing annulus 210 and a respective shape of the second wall 222 of premixing annulus 210 each is defined by an exponential function. In one example embodiment, a respective shape of the first wall 220 of premixing annulus 210 and a respective shape of the second wall 222 of premixing annulus 210 each is defined by a hockey stick function. As will be appreciated by one skilled in the art, a hockey stick mathematical function defines a curve shape that resembles the shape of an ice hockey stick, in that the shape turns from a nearly flat "blade" to a long "handle".

[0028] In operation, in disclosed embodiments the pilot burner remains concentric with the radial swirler; however, rather than being flush with the base of the radial swirler, as commonly arranged in certain known combustor designs, the body of the pilot burner extends to the point where a radial expansion of the combustion chamber is initiated. Furthermore, in contrast to certain known combustor designs, where the pilot fuel is injected radially inward into an inner recirculation zone (IRZ); in disclosed embodiments, the pilot fuel is injected radially outwardly into IRZ 226 so that the pilot fuel is more likely to reach into a root of a flame front, such as flame front 230, where the flame is stabilized in the shear layers formed by IRZ 226 and outer recirculation zone ORZ 228 and the reactants jet.

100291 Features of disclosed embodiments are believed to provide at least the following technical advantages: [0030] -bringing the pilot flame relatively closer to the main flame and thus more effectively enriching a flame front at its root, which is desirable for flame stability.

100311 -substantially reducing pilot fuel requirement to stabilize the flame front.

[0032] -avoiding or inhibiting formation of uneven flame regions, where, for example, a pilot flame and a main flame would be formed in distinct regions.

[0033] -inhibiting substantially elevated pilot burner edge (e.g., tip) temperatures due to substantial elimination of main fuel entrainment at or in the vicinity of the pilot burner edge.

[0034] -reducing the residence time of pilot fuel entrained in the IRZ.

[0035] -forming the IRZ with a relatively smaller size, thus resulting in shorter residence time, which is beneficial for inhibiting NOx emissions.

[0036] -Increasing life of nozzle guide vanes and other engine components due to lower central vortex core temperatures.

[0037] In operation, in disclosed embodiments, the pilot flame strongly interacts with the main flame which promotes static flame stability and may lead to improved dynamic stability. By way of comparison, in certain known combustor designs, a non-uniform heat release distribution is formed in the combustor due to weak interaction of the pilot flame with the main flame and this diminishes static flame stability and can lead to high dynamic instability and lean blowoff conditions; and to overcome the foregoing issues in certain known combustor designs, higher than necessary pilot fuel is typically injected in addition to fuel enrichment from the partially premixed fuel from the main fuel stream. Injection of this extra pilot fuel, however, in such known combustor designs, can result in relatively high local temperatures which promotes thermal NOx formation and high temperatures on the pilot burner face. Additionally, in such known combustor designs, the relatively large and fixed diameter of their premixing chamber increases susceptibility to flashback when burning highly reactive fuels like hydrogen, etc.; and this is because the velocity of the incoming reactants mixture is not high enough to counter the relatively higher turbulent flame speeds of the more reactive fuel mixtures. By way of comparison, features of disclosed embodiments permit imparting a relatively higher velocity to the reactants flowing in the premixing chamber and thus inhibit flashback. That is, in disclosed embodiments, the speed of the reactants flowing in the premixing chamber with a progressively reduced annulus will increase (i.e., will experience acceleration) compared to certain known premixing chambers that have a fixed unchanging diameter for their respective annuli. % [0038] Although an exemplary embodiment of the present disclosure has been described in detail, those skilled in the art will understand that various changes, substitutions, variations, and improvements disclosed herein may be made without departing from the spirit and scope of the disclosure in its broadest form.

[0039] None of the description in the present application should be read as implying that any particular element, step, act, or function is an essential element, which must be included in the claim scope: the scope of patented subject matter is defined only by the allowed claims. Moreover, none of these claims are intended to invoke a means plus function claim construction unless the exact words "means for" are followed by a participle.

[0040] Although an exemplary embodiment of the present disclosure has been described in detail, those skilled in the art will understand that various changes, substitutions, variations, and improvements disclosed herein may be made without departing from the spirit and scope of the disclosure in its broadest form.

Claims

CLAIMSWhat is claimed is: I. A combustor comprising: a radial swirler coupled to receive respective flows of reactants comprising a flow of air and a flow of main fuel; a premixing chamber defining a premixing annulus fluidly coupled to receive a swirling mix of the respective flows of reactants from the radial swirler; a pilot burner having a pilot burner body including a pilot burner edge having a plurality of pilot fuel injection orifices arranged to inject pilot fuel for combustion in a combustion chamber disposed downstream from the premixing chamber, wherein at least a portion of the pilot burner body defines a first wall of the premixing annulus, wherein a portion of a combustor liner defines a second wall of the premixing annulus, wherein the first wall and the second wall of the premixing annulus are coaxial walls configured to gradually turn from a radial direction to an axial direction, wherein the pilot burner body axially extends beyond a downstream edge of the radial swirler to a point where a radial expansion of the combustion chamber is initiated.
2. The combustor of claim 1, wherein a diameter of the premixing annulus is less than a diameter defined from side to side of the combustor liner at the point where the radial expansion of the combustion chamber is initiated.
3. The combustor of claim 1 or 2, wherein the plurality of the pilot fuel injection orifices is circumferentially disposed along a periphery of the first wall of the premixing annulus about the pilot burner edge.
4. The combustor of any one of claims 1 to 3, wherein the plurality of the pilot fuel injection orifices is arranged to radially outwardly eject the pilot fuel from the pilot burner.
5. The combustor of any one of claims 1 to 4, wherein the premixing annulus comprises a first cross sectional area at an inlet of the premixing annulus, wherein the premixing annulus comprises a second cross sectional area at an outlet of the premixing annulus, wherein a ratio of the first cross sectional area and the second cross sectional area is less than unity.
6. The combustor of claim 5, wherein the ratio of the first cross sectional area and the second cross sectional area is in a range from 0.7 to 0.9.
7. The combustor of any one of claims 1 to 5, wherein the first cross sectional area and the second cross sectional area is each respectively smaller relative to a size of a cross-sectional area defined by the combustor liner at the location where the radial expansion of the combustion chamber takes place
8. The combustor of claim 1, wherein a respective shape of the first wall and a respective shape of the second wall of the premixing annulus each is defined by a parabolic function.
9. The combustor of claim 1, wherein a respective shape of the first wall and a respective shape of the second wall of the premixing annulus each is defined by a polynomial function.
10. The combustor of claim 1, wherein a respective shape of the first wall and a respective shape of the second wall of the premixing annulus each is defined by an exponential function.
11. The combustor of claim 1, wherein a respective shape of the first wall and a respective shape of the second wall of the premixing annulus each is defined by a hockey stick mathematical function.