EP0813668B1 - Verfahren zur verbrennung eines brennstoffs in einer gasturbine sowie entsprechende gasturbine - Google Patents

Verfahren zur verbrennung eines brennstoffs in einer gasturbine sowie entsprechende gasturbine Download PDF

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Publication number
EP0813668B1
EP0813668B1 EP96905682A EP96905682A EP0813668B1 EP 0813668 B1 EP0813668 B1 EP 0813668B1 EP 96905682 A EP96905682 A EP 96905682A EP 96905682 A EP96905682 A EP 96905682A EP 0813668 B1 EP0813668 B1 EP 0813668B1
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EP
European Patent Office
Prior art keywords
turbine
flow
gas turbine
fuel
compressor section
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Expired - Lifetime
Application number
EP96905682A
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German (de)
English (en)
French (fr)
Other versions
EP0813668A1 (de
Inventor
Manfred Ziegner
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Siemens AG
Siemens Corp
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Siemens AG
Siemens Corp
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Publication date
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Publication of EP0813668A1 publication Critical patent/EP0813668A1/de
Application granted granted Critical
Publication of EP0813668B1 publication Critical patent/EP0813668B1/de
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
    • F23R3/425—Combustion chambers comprising a tangential or helicoidal arrangement of the flame tubes

Definitions

  • the invention relates to a method for the combustion of a Fuel in a stream of compressed air, which is a Gas turbine from a compressor part to a turbine part flows through, the fuel the current in the compressor part fed and between the compressor part and the turbine part is burned.
  • the invention also relates to a corresponding one Gas turbine.
  • the Gas turbine as a "gas turbine group”
  • the compressor part as a “compressor”
  • the turbine part referred to as “turbine”.
  • the different designations are due to one inconsistent use of the term "gas turbine” in the Experts.
  • both a turbine can actually Senses, that is, a power machine made up of one Electricity heated gas extracted mechanical energy when also a unit from a turbine in the real sense called combustion chamber or combustion chambers and compressor part become.
  • a "Gas turbine” always understood a unit that next to one Turbine in the actual sense, always as a “turbine part” referred to, at least one associated compressor part includes.
  • burners that can be used in a gas turbine are EP 0 193 838 B1, US Patent Re. 33896, the EP 0 276 696 B1 and US Pat. No. 5,062,792.
  • a Combustion chamber in the form of an annular combustion chamber with a large number of torches arranged in a ring is in the EP 0 489 193 A1.
  • thermodynamic losses An important source of thermodynamic losses is that between the compressor part and the turbine part, ie over the area of the gas turbine where the electricity is more compressed Air is heated by burning a fuel, occurring Pressure loss caused by the so far always high construction effort to implement a combustion device in the form of one or more Combustion chambers.
  • Certain approaches to reduce effort are known; particularly goes from the already mentioned EP 0 590 297 A1 discloses a so-called "ring combustion chamber", in the current has a swirl that impressed it in the compressor part is maintained during the combustion of the fuel Should, so that there is a conventional at the inlet of the turbine part fixed blade ring with which to operate of the turbine part required swirl built up should be left out.
  • U.S. Patent 2,630,678 according to which the feeding of Fuel can already take place in the compressor part.
  • Turbine inlet temperature that is the temperature of the Electricity after combustion of the fuel and when entering the turbine part.
  • Turbine inlet temperature is limited through the resilience of the components of the turbine part, which is given in particular by the resilience of the materials used and any provided Measures to cool the components. Find such measures usually its limit is that necessary for cooling Air must be drawn from the stream and not more is available for combustion. Also from The distribution of the temperature in the stream is important Entry into the turbine section.
  • the cause of the inhomogeneous distribution of the temperature in a current in a conventional gas turbine is usually in the complex and inherently inhomogeneous treatment of electricity and fuel between the compressor part and the turbine part. This applies in particular if the flow is divided into partial flows and several combustion chambers or several individual burners; this applies also in conventional ring combustion chambers, each with a To a large extent do without subdivision of the current, but always several more and necessarily spaced apart Provide burners to heat the electricity.
  • the guide device in the turbine part is thermal most heavily loaded component and must be correspondingly complex be executed; in addition, this already takes place Guide device a partial relaxation, and thus temperature reduction, of the flue gas in the stream. Accordingly the first rotating turbine stage does not determine the maximum possible temperature of the current, but the control device at the inlet of the turbine part, which, however, does not flow Energy extracted.
  • the invention is based on Task, a method and a gas turbine of the aforementioned Specify the type of combustion of the fuel in the stream while ensuring as uniform as possible Distribution of the temperature in the stream and avoiding it of losses.
  • this task is solved specified a method for burning a fuel in a stream of compressed air which is a gas turbine from flows through a compressor part to a turbine part, wherein the fuel is supplied to the flow in the compressor section and between the compressor part and the turbine part is burned, where the current exits the compressor part first, with an adjustable last idler at the outlet end adjustable flow direction impressed on the compressor part which is between the compressor part and the turbine part is transformed into a second flow direction, the one Corresponds to the nominal flow direction with which the current flows directly fed to a first impeller of the turbine part must be, the first flow direction depending is regulated by the operating state of the gas turbine.
  • the current is a first when exiting the compressor part Directed flow direction, which is due to the combustion of the fuel in the stream in a second flow direction is transformed, that of a nominal flow direction corresponds for which the turbine part is designed.
  • a Heating especially when burning the Fuel occurs, the flow direction changes.
  • the heating causes an increase in speed, with which the current propagates, although only a component of the speed in the direction of propagation of the stream increases.
  • the component representing the swirl the speed can be perpendicular to the direction of propagation naturally not changed by the heating of the current become.
  • the first flow direction with which the current emerges from the compressor part, that the second flow direction that the current at Has entry into the turbine part, with one through the geometry direction of the turbine section, here "nominal flow direction" called lies.
  • the first Flow direction that is the flow direction with which the current exits the compressor section, depending of a thermal power with which by combustion Heat is generated to regulate. It is understood that a regulation depending on the thermal performance in the result also a regulation depending on one of the gas turbine mechanical power output.
  • the compressor part includes a last idler, which of the Current flows through as it emerges from the compressor part, and which to vary the first flow direction with which the current flows behind the last stator, adjustable is.
  • Adjustable guide wheels for compressor parts are essential known, however, according to previous practice only used at the inlet of a compressor part and serve to adjust the inlet cross-section through which Air is sucked in.
  • the adjustable Stator in particular to adjust the performance by the gas turbine is to be delivered.
  • With an adjustable last idler at the outlet end of a compressor part can the twist with which the current leaves the compressor part is set be, and this in particular depending on the Operating state of the gas turbine. This makes it possible to Swirl of the current for every conceivable operating condition the requirements, the turbine part to the swirl of the stream poses to adapt.
  • pilot flames directed into the stream in reasonable Number provided can be used with small Burners are formed. They cause local heating and ignition of the fuel-air mixture, which can can spread quickly across the entire stream.
  • the stream after the mixing is delayed with the fuel.
  • a delay which in particular in a trained as a diffuser Annular channel between the compressor part and the turbine part can be done for stable combustion set the favorable speed of the current. Possibly this delay can also be fixed in a special Blade ring made; on such a blade ring can also use stabilization devices suitable for combustion.
  • the method is preferably used using a fuel in the form of a combustible gas, in particular Natural gas or coal gas, with any of coal gas combustible gaseous product of a coal gasification process is to be understood.
  • a fuel in the form of a combustible gas, in particular Natural gas or coal gas, with any of coal gas combustible gaseous product of a coal gasification process is to be understood.
  • the solution to the problem Specified a gas turbine for combustion of a task Fuel in a stream of compressed air, which of a compressor part flows to a turbine part, with a Ring channel for guiding the flow and nozzles for feeding of the fuel to the stream in the compressor part at which Gas turbine the current when exiting the compressor part a first, with an adjustable last idler on Outlet end of the compressor part adjustable flow direction is stamped between the compressor part and turbine part is transformed in a second flow direction, which corresponds to a nominal flow direction with which the Current supplied directly to a first impeller of the turbine part must be, the first flow direction depending is regulated by the operating state of the gas turbine.
  • the nozzles are preferably on a stator in the compressor part attached and can be fixed in particular Guide vanes, the essential components of the guide wheel are integrated.
  • the nozzles are preferably in hollow guide vanes of the guide wheel attached.
  • the idler with the nozzles is the penultimate or last one, by the flow around the stator, in connection with a Preferred training described below, the penultimate diffuser.
  • Such placement of the nozzles ensures with even distribution of the fuel in the A high security against premature ignition of the current Fuel, as in view of that in a more modern one Gas turbine at the compressor outlet temperature prevailing is.
  • a flame holder arranged between the compressor part and the turbine part.
  • a flame holder is, for example, as Flow obstacle trained and causes in the Current immediately following the flame holder or backflow area.
  • Such a vortex is suitable for the formation of a largely stationary flame, what to ensure stable and complete Combustion can be important.
  • the annular channel between the compressor part is also preferred and expanded the turbine part like a diffuser.
  • This expansion does not necessarily have to be even, but may be more or less erratic his. This leads to the formation of a front in the Electricity, at which it slows down significantly and at which can form and maintain a stable flame; so he can Diffuser act as a flame holder.
  • the ring channel between the compressor part is further preferred and the turbine part with ceramic heat shield elements lined which is the originating from the combustion Take up thermal load with low cooling requirements.
  • the gas turbine also has a turbine part, in which the current is fed directly to an impeller becomes. This implies that the current in the ring channel is performed with a twist and that the combustion in this Electricity takes place.
  • the turbine part is particularly simple designed since it does not require a stator at its inlet, with one for operating the rotating impellers of the turbine part required swirl would have to be built up first.
  • a stator is at the inlet of the turbine part one of the most thermally stressed components the gas turbine with a correspondingly high cooling requirement, the conventionally at the expense of those available for combustion Air must be covered, as well as with corresponding requirements the material to be used for the production.
  • Using the invention can therefore be a particularly economical Gas turbine can be realized.
  • the figure shows a gas turbine 1 with a compressor part 2 and a turbine part 3.
  • the compressor part 2 which is only partially is shown, sucks from the surroundings of the gas turbine 1 air, compresses it and makes it more compressed as stream 4 Air ready.
  • This stream 4 is in the compressor part 2 mixed with fuel 5, this fuel 5 by Nozzles 6 is supplied.
  • the stream 4 has a first flow direction 7, with a velocity component that is perpendicular to the direction in which the stream 4 propagates, is directed.
  • This first flow direction 7 changes under certain circumstances, until the current 4 reaches the turbine part 3, whereby a second flow direction at the inlet of the turbine part 3 8 sets.
  • the change is essentially due to the Combustion of the fuel 5, which is initiated by pilot flames 9 between the compressor part 2 and the turbine part 3 protrude into stream 4.
  • the pilot flames 9 form on fuel, which through appropriate nozzles 10th is delivered.
  • At the entrance of the turbine part 3 is not a fixed idler according to conventional practice, but immediately an impeller 11.
  • an impeller 11 By setting the second flow direction accordingly 8 can namely on a stator at the inlet of the turbine part 3 can be dispensed with.
  • the fuel 5 is supplied to the stream 4 by the mentioned nozzles 6, which are on a penultimate stator 12 of the compressor part 2 are.
  • the nozzles 6 are special Mouths of channels in corresponding hollow guide vanes, arranged together and in a ring the penultimate Form stator 12.
  • a last guide wheel 13 is arranged, which consists of guide vanes is formed with corresponding adjustment devices 14 are adjustable.
  • the gas turbine 1 the first flow direction 7 and thus the second flow direction 8 set and in particular the Requirements of the turbine part 3 can be adapted.
  • the gas turbine 1 can possibly be from a stator 12 are disregarded at the outlet of the compressor part 2.
  • the compressor part 2 In order to stabilize the combustion of the fuel 5 in the stream 4, are between the compressor part 2 and the turbine part 3 flame holder 15 provided.
  • this flame holder 15 is less popular, especially since Flame holders are known in many ways in the prior art are and can be used in the present case.
  • the embodiment is the flame holder 15 for Example a firmly anchored rod, which is in an annular channel 16, through which the flow 4 from the compressor part 2 to the turbine part 3 continues, protrudes. It is important that behind the flame holder 15 forms a vortex on which can stabilize a flame. This function can not only of bars, but also of differently designed components be perceived.
  • the fuel 5 is supplied to the nozzles 6 and 10 via corresponding fuel lines 17 and fuel pumps 18 from a fuel supply 19.
  • the fuel supply 19 can be any memory, but it is also conceivable that the fuel supply 19 is a public supply network, especially for gaseous fuel such as natural gas. Also it is conceivable that the fuel supply 19 belongs to a system, gasified in the coal and a combustible gasification product, namely coal gas, which is used as fuel the gas turbine 1 can serve.
  • heat shields are of various types in the relevant state of the art known, so that further explanations are unnecessary at this point are.
  • the invention relates to a gas turbine and a method for the combustion of a fuel compressed in a stream Air that takes a gas turbine from one compressor part to another Flows through the turbine part, the fuel between the compressor part and the turbine part is burned, wherein the fuel is supplied to the current in the compressor section.
  • the invention enables a significant simplification of the Building a gas turbine and bringing by avoiding Pressure losses and friction losses also have significant advantages in terms of the thermodynamics of the in the gas turbine energy conversion process taking place.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP96905682A 1995-03-06 1996-03-05 Verfahren zur verbrennung eines brennstoffs in einer gasturbine sowie entsprechende gasturbine Expired - Lifetime EP0813668B1 (de)

Applications Claiming Priority (3)

Application Number Priority Date Filing Date Title
DE19507763A DE19507763A1 (de) 1995-03-06 1995-03-06 Verfahren und Vorrichtung zur Verbrennung eines Brennstoffs in einer Gasturbine
DE19507763 1995-03-06
PCT/DE1996/000386 WO1996027764A1 (de) 1995-03-06 1996-03-05 Verfahren zur verbrennung eines brennstoffs in einer gasturbine sowie entsprechende gasturbine

Publications (2)

Publication Number Publication Date
EP0813668A1 EP0813668A1 (de) 1997-12-29
EP0813668B1 true EP0813668B1 (de) 2001-07-25

Family

ID=7755753

Family Applications (1)

Application Number Title Priority Date Filing Date
EP96905682A Expired - Lifetime EP0813668B1 (de) 1995-03-06 1996-03-05 Verfahren zur verbrennung eines brennstoffs in einer gasturbine sowie entsprechende gasturbine

Country Status (7)

Country Link
US (1) US6003297A (enrdf_load_stackoverflow)
EP (1) EP0813668B1 (enrdf_load_stackoverflow)
JP (1) JP3939753B2 (enrdf_load_stackoverflow)
DE (2) DE19507763A1 (enrdf_load_stackoverflow)
ES (1) ES2160804T3 (enrdf_load_stackoverflow)
IN (1) IN187803B (enrdf_load_stackoverflow)
WO (1) WO1996027764A1 (enrdf_load_stackoverflow)

Families Citing this family (12)

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DE19641725A1 (de) * 1996-10-10 1998-04-16 Asea Brown Boveri Gasturbine mit einer sequentiellen Verbrennung
US6192668B1 (en) * 1999-10-19 2001-02-27 Capstone Turbine Corporation Method and apparatus for compressing gaseous fuel in a turbine engine
US20020166324A1 (en) 1998-04-02 2002-11-14 Capstone Turbine Corporation Integrated turbine power generation system having low pressure supplemental catalytic reactor
DE10035676A1 (de) * 2000-07-21 2002-02-07 Siemens Ag Gasturbine und Verfahren zum Betrieb einer Gasturbine
EP1780387A3 (en) * 2000-09-05 2007-07-18 Sudarshan Paul Dev Nested core gas turbine engine
US7784261B2 (en) * 2006-05-25 2010-08-31 Siemens Energy, Inc. Combined cycle power plant
GB0617925D0 (en) * 2006-09-12 2006-10-18 Rolls Royce Plc Components for a gas turbine engine
US8387389B2 (en) * 2007-12-20 2013-03-05 Volvo Aero Corporation Gas turbine engine
US8006500B1 (en) * 2008-01-29 2011-08-30 Florida Turbine Technologies, Inc. Swirl combustor with counter swirl fuel slinger
US20150000298A1 (en) * 2013-03-15 2015-01-01 Advanced Green Technologies, Llc Fuel conditioner, combustor and gas turbine improvements
US9599019B2 (en) * 2014-02-05 2017-03-21 United Technologies Corporation Dual oil supply tube
US11499485B2 (en) * 2020-02-10 2022-11-15 Raytheon Technologies Corporation Engine control device and methods thereof

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Also Published As

Publication number Publication date
JP3939753B2 (ja) 2007-07-04
EP0813668A1 (de) 1997-12-29
JPH11501380A (ja) 1999-02-02
ES2160804T3 (es) 2001-11-16
US6003297A (en) 1999-12-21
DE19507763A1 (de) 1996-09-12
IN187803B (enrdf_load_stackoverflow) 2002-06-29
WO1996027764A1 (de) 1996-09-12
DE59607363D1 (de) 2001-08-30

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