Gas turbine arrangement with controlled bleed air injection into combustor, and method of operation

Abstract

A gas turbine arrangement for dual fuel operation has a first manifold that delivers a first fuel or compressor bleed fluid and is connected to a bleed port and a first passage for ejecting fuel or fluid into a combustor space. A second manifold delivers a second fuel and is connected to a second passage for ejecting the second fuel into the combustor space. A control system, when operated with the second fuel, provides the second fuel to the second manifold and continuously opens the bleed valve to provide bleed fluid into the first manifold to replace the first fuel. The control system controls the bleed valve over time by throttling a mass flow of the bleed fluid provided to the first passage or by increasing a mass flow of the bleed fluid provided to the first passage to adapt to fuel properties of the second fuel.

Claims

1. A gas turbine arrangement for at least dual fuel operation, comprising: a compressor; a bleed port formed in one of the compressor or a transition duct; a bleed valve in fluid connection with the bleed port; a plurality of combustors, each combustor of the plurality of combustors having a combustor space; a first manifold connected to the each combustor and the bleed valve; a second manifold connected to the each combustor and the bleed valve; a first further bleed valve connected to the bleed valve and the first manifold: a second further bleed valve connected to the bleed valve and the second manifold; and a control system; wherein the first manifold is configured to deliver either a first fuel or a compressor bleed fluid, the first manifold being connectable to the bleed port via a bleed valve for delivering the compressor bleed fluid to the each combustor, wherein the first manifold is connected to at least one first passage of the each combustor for ejecting the first fuel or the compressor bleed fluid into the combustor space; wherein the second manifold is configured to deliver a second fuel different from the first fuel, wherein the second manifold is connected to at least one second passage of the each combustor for ejecting the second fuel into the combustor space; wherein the control system, when the each combustor is operated solely with the second fuel, is configured to continuously open the bleed valve and the first further bleed valve to provide the compressor bleed fluid into the first manifold and to close the second further bleed valve to provide the second fuel into the second manifold to monitor a flame behaviour and/or a fuel distribution in the combustor space, and to modify the flame behaviour and/or the fuel distribution by controlling the bleed valve over time by throttling a mass flow of the compressor bleed fluid provided to the at least one first passage or alternatively by increasing a mass flow of the compressor bleed fluid provided to the at least one first passage.

2. The gas turbine arrangement according to claim 1, wherein the control system is configured, when the each combustor is operated solely with the second fuel, to control additionally the first further bleed valve and/or further fuel supply valves over time by throttling or alternatively increasing a mass flow of the compressor bleed fluid provided to the at least one first passage.

3. The gas turbine arrangement according to claim 2, wherein the control system is configured, for the monitoring of the flame behaviour and/or fuel distribution, to monitor an air to fuel ratio in the combustor space and/or an axial velocity of a premixed air and fuel mixture in the combustor space, wherein an axial direction is defined as a direction of a main expanse of the combustor space.

4. The gas turbine arrangement according to claim 1, wherein the control system is configured, for control of the bleed valve, to evaluate fuel properties of the second fuel.

5. The gas turbine arrangement according to claim 4, wherein the fuel properties of the second fuel are related to a heating value and/or a chemical composition of the second fuel.

6. The gas turbine arrangement according to claim 1, wherein the control system is configured, when the each combustor is operated solely with the first fuel, to provide the first fuel to the first manifold and to the second manifold.

7. The gas turbine arrangement according to claim 1, wherein the compressor bleed fluid is extracted from one of trailing stages of the compressor.

8. The gas turbine arrangement according to claim 7, wherein the compressor bleed fluid is extracted from a last stage of the compressor, or from a transition duct downstream of the last stage of the compressor.

9. The gas turbine arrangement according to claim 1, wherein the first manifold is configured to provide the first fuel as main fuel of the each combustor to a first set of main fuel nozzles, and wherein the second manifold is configured to provide at least the second fuel as an alternative main fuel of the each combustor to a second set of the main fuel nozzles.

10. The gas turbine arrangement according to claim 9, wherein the first set of the main fuel nozzles have different aperture size and/or are positioned at different locations compared to the second set of the main fuel nozzles.

11. The gas turbine arrangement according to claim 9, wherein the second manifold is configured to provide also the first fuel as a main fuel of the each combustor when the each combustor is operated solely with the first fuel.

12. The gas turbine arrangement according to claim 1, further comprising: a cooling device for cooling the compressor bleed fluid.

13. The gas turbine arrangement according to claim 12, wherein the cooling device is controlled by the control system, to adjust the cooling of the compressor bleed fluid.

14. The gas turbine arrangement according to claim 1, wherein the first fuel and/or the second fuel is gaseous.

15. A method to operate a gas turbine arrangement, the method comprising: operating a compressor to provide a compressor fluid; delivering the compressor fluid to a plurality of combustors, each combustor of the plurality of combustors having a combustor space; delivering a first fuel into a first manifold; delivering a second fuel different from the first fuel into a second manifold; closing a second further bleed valve connected to a bleed valve to provide the second fuel to the second manifold when the each combustor is operated solely with the second fuel; continuously opening the bleed valve and a first further bleed valve connected to the a bleed valve, to provide a portion of the compressor fluid through a bleed port formed in one of the compressor or a transition duct as a compressor bleed fluid into the first manifold; monitoring a flame behaviour and/or a fuel distribution in the combustor space; and controlling the bleed valve over time by one of throttling a mass flow of the compressor bleed fluid provided to at least one first passage, and increasing a mass flow of the compressor bleed fluid provided to the at least one first passage to adapt to fuel properties of the second fuel for modifying the flame behaviour and/or the fuel distribution.

16. The method to operate a gas turbine arrangement according to claim 15, further comprising: providing a high reactive fuel as the second fuel to the second manifold, the high reactive fuel being more reactive than the first fuel, or more reactive than natural gas.

17. The method to operate a gas turbine arrangement according to claim 16, wherein the high reactive fuel is rich of hydrogen, ethane, propane, and/or hydrocarbons heavier than ethane and propane.

18. The method to operate a gas turbine arrangement according to claim 15, wherein the monitoring of the flame behaviour comprises: monitoring of combustion pressure dynamics, and/or evaluating pressure sensor data, and/or monitoring of flashbacks, and/or evaluating thermocouple sensor data.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

(1) Embodiments of the invention will now be described, by way of example only, with reference to the accompanying drawings, of which:

(2) FIG. 1: shows schematically of an abstract configuration of an exemplary gas turbine arrangement;

(3) FIG. 2: shows a cross sectional but still simplified view of one embodiment of the gas turbine arrangement;

(4) FIG. 3: shows an exemplary burner used within the gas turbine arrangement.

(5) The illustration in the drawing is schematic. It is noted that for similar or identical elements in different figures, the same reference signs will be used.

(6) Some of the features and especially the advantages will be explained for an assembled and operating gas turbine, but obviously the features can be applied also to the single components of the gas turbine but may show the advantages only once assembled and during operation. But when explained by means of a gas turbine during operation none of the details should be limited to a gas turbine while in operation.

DETAILED DESCRIPTION OF THE INVENTION

(7) In reference to FIG. 1 and FIG. 2, it is briefly explained the principles of a gas turbine engine which comprises the features of a gas turbine arrangement 1 as introduced before.

(8) In respect of operating a gas turbine engine, a gas turbine—short for gas-turbine engine—comprises an air inlet or intake 10 at one axial end followed by a compressor 11 in which incoming air (indicated in FIG. 1 by an arrow) is compressed for application to one or more combustors 12, which may be annular, so-called can-annular or of silo type, the can-annular type being distributed circumferentially around a turbine axis defining a rotational axis 13. Fuel is introduced into the combustors 12 and therein mixed with a major part of the compressed air taken from the compressor 11. Hot gases with high velocity as a consequence of combustion in the combustors 12 are directed to a set of turbine blades within a turbine section 16, being guided (i.e. redirected) by a set of guide vanes. The turbine blades and a shaft—the turbine blades being fixed to that shaft—form a part of the rotor and are rotated about the rotational axis 13 as a result of the impact of the flow of the hot gases. That rotating rotor (or in case of a plurality of rotors another connected rotor) also rotates blades of the compressor 11, so that the compressed air supply to the combustors 12 is generated by the rotor blades of the compressor 11 (i.e. by the used rotating blades interacting with static vanes) once in operation. There may be more than one rotor and/or more than one rotor shaft in the gas turbine engine. Past the turbine section 16, at a second axial end of the gas turbine and downstream of the turbine section 16, the combusted gas will be guided to an exhaust 14 of the gas turbine via which it will be released in to the ambient air or guided to a further component, like a steam boiler or a steam turbine. The exhausted gas may contain also combustion emissions, like CO.

(9) In the following we will identify the ambient air that is guided from the air inlet, via the compressor 11 and to the combustor 12 as “compressor fluid”, guided through the mentioned components via a main fluid path, which is an annular path via which the compressor fluid is delivered and in which the physical properties of the compressor fluid will be changed. The compressor fluid is delivered via the main fluid path to the combustor 12 for combustion with fuel.

(10) Note that in the following a location of one component in relation to a second one is made in respect of a flow direction of a fluid, usually the compressor fluid or in other cases of fuel. The term “upstream” defines a position that the fluid passes earlier than the “downstream” position. That means an upstream end of a compressor is the section at which air enters the compressor from the inlet of the gas turbine. A downstream end defines the position at which the air is discharged from the compressor. The direction from an upstream end of the compressor to a downstream end of the compressor also is defined as (positive) axial direction. A direction perpendicular to a rotational axis of the gas turbine engine is called radial direction. Furthermore a direction given from a location on a cylindrical surface may comprise an axial vector component, a radial vector component and a circumferential vector component, all three components being perpendicular to another. Thus, a circumferential direction is the direction lateral to a cylindrical surface or lateral to an annular cavity.

(11) FIG. 1 shows in an abstract way a bleed port 30 located at a downstream end of the compressor 11 or at a transition duct (identified in FIG. 2 via reference numeral 18) From the bleed port 30 compressor bleed fluid 31 is diverted from the main fluid path into a passage system, including several valves. A bleed valve 32 is present to close or open the bleed port 30 fluidically. In the exemplary configuration the shown gas turbine arrangement 1 the passage system expands to three lines with three further bleed valves 32′, downstream of the bleed valve 32, allowing compressor bleed fluid 31 to be injected into a plurality of fuel manifolds, i.e. a first manifold 21, a second manifold 22, and a third manifold 23. A different number of manifolds may be present.

(12) The same plurality of fuel manifolds are also connected to a fuel supply 15, which advantageously is configured to provide at least two different fuel types. Further fuel supply valves 43, i.e. a first fuel valve 43′, a second fuel valve 43″, and a third fuel valve 43′″ may be present and allow connection to the previously mentioned fuel manifolds 21,22,23.

(13) The exemplary configuration should allow individual supply of fuel or air to the different manifolds. This can be achieved also by alternative configurations. Furthermore the valves are advantageously valves that can open and close the throughput, but are also able to throttle or to expand the throughput.

(14) The three manifolds 21,22,23 are connected to each of the combustors 12. In FIG. 1 two combustors are shown as a pure example for explaining the invention.

(15) Based on the basic setup of FIG. 1, a more detailed view is shown in FIG. 2 in respect of a control system 17 which is configured to control all the different valves.

(16) In FIG. 2 the bleed port 30 is shown to divert from one of the last stages of the compressor 11. It is assumed that the pressure of the compressor bleed fluid 31 is sufficiently high to purge at least one of the fuel manifolds 21,22,23, depending on the mode of operation. In FIG. 2 two fuel manifolds 21 and 22 are shown. The first manifold 21 is connected to a first fuel valve 43′ and a second fuel valve 43″. The second manifold 22 is connected to a fourth fuel valve 43″″. A first fuel—advantageously natural gas—can be provided via the first fuel valve 43′ and via the fourth fuel valve 43″″ to both manifolds 21 and 22.

(17) Compressor bleed fluid as extracted from bleed port 30 can be provided via bleed valves 32—two bleed valves are indicated in FIG. 2, one per manifold—to both fuel manifolds 21,22. In this configuration one valve 32 is shown per manifold 21,22.

(18) All these valves are configurable via control lines 60 from the control system 17.

(19) In this example the control system 17 is configured to provide less reactive fuel to the first manifold 21 and also to the second manifold 22. Further the control system 17 is configured to provide higher reactive fuel to—only—the second manifold 22. In parallel, when both manifolds 21,22 are provided with fuel, the bleed valves 32 are closed by the control system 17. In case when only the second manifold 22 is provided with fuel, the bleed valves 32 will be set to provide compressor bleed fluid to—only—the first manifold 21, allowing to purge the first manifold 21.

(20) The advantages may not become apparent in FIG. 2 alone, but based on FIG. 2, a burner configuration is explained in more detail in relation to FIG. 3, so that also the advantages of this configuration should become apparent.

(21) In FIG. 3 an exemplary combustor 12 is shown in a sectional view. This is only one combustor 12 of possibly a plurality of combustors. The combustor 12 may be substantially rotational symmetrical. A downstream direction within the combustor is indicated by an arrow for an axial direction A. Sections of the manifolds 21,22 are shown, which branch off to individual passages for supplying fuel nozzles within the combustor 12. The first manifold 21 is connected via a first passage 41 to a first set of main fuel nozzles 71. The second manifold 22 is connected via a second passage 42 to a second set of main fuel nozzles 72.

(22) The fuel nozzles are located in the given example on a surface of a swirler vane of a swirler for mixing the provided fuel via the nozzles with compressor fluid 33, the latter entering swirler passages through the swirler. The swirler is part of a burner 57.

(23) The combustor 12 comprises the burner 57 and a combustor space 50. The combustor space 50 starts upstream at the swirler passages, continues later via a premixing space 52 and ends in the combustion chamber 51.

(24) A fuel and air mixture travels through the premixing space 52 until ignition in the combustion chamber 51. An exemplary flame front 53 is depicted as a pure fictitious example. An axial velocity 54 of the fuel and air mixture through the premixing space 52 is shown as arrows in axial direction A.

(25) As previously indicated in relation to FIG. 2, one mode of operation may be supply of natural gas (as the first fuel) via both of the manifolds 21,21, which results in provision of fuel via both the first set of main fuel nozzles 71 and the second set of main fuel nozzles 72.

(26) In another mode of operation with a highly reactive fuel as the second fuel, the control system 17 is configured to only provide this second fuel via the second manifold 22 and in consequence only via the second set of main fuel nozzles 72. Furthermore, the bleed valve 32 (and/or one of the further bleed valves 32′ if present) is opened so that the first manifold 21 is purged with air or compressor bleed fluid 31. In consequence this compressor bleed fluid 31 is exhausted via the first set of main fuel nozzles 71. This has the advantage that the fuel to air ratio is reduced when operated with the second fuel compared to operation with the first fuel. A second advantage occurs as air jets are exhausted through the first set of main fuel nozzles 71 which possibly increases mixing of fuel and air. As an advantage, the flame front 53 may substantially keep position and is not drawn into upstream direction within the premixing space 52.

(27) The amount of air provided through the first set of fuel nozzles 71 may be above a threshold below which a fuel and air mixture would enter backwards into the first set of main fuel nozzles 71. The amount of air can be controlled by the control unit 17.

(28) The control unit 17 is configured to monitor a flame behaviour and/or fuel distribution in the combustor space 50. The flame behaviour may be derived from data collected from at least one pressure sensor 55 and from at least one thermocouple sensor 56. Possibly a plurality of these sensors may be present around a circumference of the combustion space 50.

(29) Based on the collected data all involved valves—particularly the bleed valve 32 (and the further bleed valves 32′ if present)—may be controlled by the control unit 17. This allows modifying the flame behaviour and/or the fuel distribution within the combustion space 50.

(30) The control is performed over time and over space. Changes of the flame behaviour and/or the fuel distribution over time can be detected and such that modifications in the fuel and air mixture can be initiated.

(31) Uneven flame behaviour and/or fuel distribution in respect of its location—i.e. over space—can also be detected and possibly be improved by changing valve settings.

(32) Advantageously combustion pressure dynamics are monitored, particularly by evaluating pressure sensor data. Additionally flashbacks—i.e. travel of the flame in negative axial direction for at least a short time—may be monitored, particularly comprising evaluating thermocouple sensor data.

(33) Different configurations can be foreseen. The positions and numbers of the mentioned valves are purely exemplary. Valves may also be replaced by other components that have the effect of restricting a mass flow through passages.

(34) The invention is particularly advantageous if two different gaseous fuels are used alternating in the same gas turbine arrangement. As a gas turbine engine is typically optimised for operation with one specific fuel type—typically natural gas—operation with a different fuel can be improved with the invention. Especially when the second fuel is a highly reactive fuel—the second fuel is rich of hydrogen, ethane, propane, and/or hydrocarbons heavier than ethane and propane, thus including butanes, pentanes, hexanes, etc.—this invention allows stable operation, as more air is provided to the combustor and the fuel and air mixture will have a different percentage than when operated with natural gas.

(35) As a brief summary, the invention is advantageously directed to adapt to fuel properties of a provided fuel and to allow operating the combustor with a variety of different fuels without the need to replace burner parts manually. An amount of provided bleed or purge fluid is determined and selected permanently or enduringly based on the flame behaviour and/or fuel distribution for the provided fuel. While the provided fuel is provided via one fuel manifold and to corresponding fuel injectors, the bleed or purge air is provided to another fuel manifold and to related fuel injectors (which are purged then). So for highly reactive fuels the bleed or purge air may be increased permanently to create a leaner fuel and air mix. The flame position may change as well in the combustion space. This allows avoiding overly hot machine parts and unwanted flame behaviour.