EP1507120A1 - Turbine à gaz - Google Patents

Turbine à gaz Download PDF

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Publication number
EP1507120A1
EP1507120A1 EP03018412A EP03018412A EP1507120A1 EP 1507120 A1 EP1507120 A1 EP 1507120A1 EP 03018412 A EP03018412 A EP 03018412A EP 03018412 A EP03018412 A EP 03018412A EP 1507120 A1 EP1507120 A1 EP 1507120A1
Authority
EP
European Patent Office
Prior art keywords
combustion air
swirl
combustion
fuel
gas turbine
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.)
Withdrawn
Application number
EP03018412A
Other languages
German (de)
English (en)
Inventor
Malte Dr. Blomeyer
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
Original Assignee
Siemens AG
Siemens Corp
Priority date (The priority date 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 date listed.)
Filing date
Publication date
Application filed by Siemens AG, Siemens Corp filed Critical Siemens AG
Priority to EP03018412A priority Critical patent/EP1507120A1/fr
Publication of EP1507120A1 publication Critical patent/EP1507120A1/fr
Withdrawn legal-status Critical Current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
    • F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
    • F23R3/50—Combustion chambers comprising an annular flame tube within an annular casing
    • 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
    • F23R3/04—Air inlet arrangements
    • F23R3/10—Air inlet arrangements for primary air
    • F23R3/12—Air inlet arrangements for primary air inducing a vortex
    • F23R3/14—Air inlet arrangements for primary air inducing a vortex by using swirl vanes
    • 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
    • F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
    • F23R3/34—Feeding into different combustion zones
    • F23R3/343—Pilot flames, i.e. fuel nozzles or injectors using only a very small proportion of the total fuel to insure continuous combustion

Definitions

  • the invention relates to an axial gas turbine, comprising a Compressor, an annular combustion chamber and a turbine part.
  • combustion instabilities can also be used to build up so-called combustion oscillations lead, where caused by the instabilities Pressure pulses from the combustion chamber wall to the flame zone reflected be there and periodically reinforce the instabilities, which in a positive feedback to a stable Burning vibration can lead.
  • combustion vibrations are both in terms of acoustic emissions as well as in terms of damaging mechanical Vibrations undesirable. Often, therefore, is the lean Vormischverbrennung stabilized by additional diffusion burners.
  • a typical burner arrangement is a ring-channel premix burner, the one central diffusion or pilot burner surrounds.
  • WO 02/08592 shows a gas turbine.
  • the gas turbine has one annular combustion chamber, short annular combustion chamber, which around the Turbine axis is arranged around.
  • a number of premix burners protrude along the circumference which are stabilized by central pilot burners.
  • a particularly efficient combustion process results here by an admixture of fuel already at the compressor outlet, whereby combustion air and fuel especially be well mixed. Branched off from the combustion air flow Cooling air for turbine blades in the turbine part is thereby also already interspersed with fuel, which when flowing out this cooling air from openings in the turbine blades leads to a reheating in the turbine part. This has one particularly high efficiency.
  • EP 590 297 shows a gas turbine group.
  • the swirling combustion air flow from the To direct compressor as directly as possible to the turbine part to there exploit the swirl so that a first row of vanes can be saved.
  • the swirling the residence time of the fuel mixed Combustion air in the combustion chamber as high as possible to for a short size, sufficient residence time for the burnout to reach in the combustion chamber. This is done by achieved that the compressor directly into the combustion chamber opens, with a last Ver Whyrleitschaufelsch also if necessary can be saved or only one Twist reinforcement is needed.
  • the US-PS 6,003,297 shows a gas turbine, in which also the compressor opens directly into the combustion chamber.
  • main fuel flow is already in the compressor added, so that despite the immediate confluence of the Combustion air into the combustion chamber before to a good mixing of combustion air and fuel comes.
  • Additional pilot burner opening into the combustion chamber stabilize the combustion. Again, through maintaining the swirl from the compressor a first Leitschaufelsch be saved in the turbine part.
  • the object of the invention is the specification of a gas turbine, in the at a high combustion stability, a particularly low Pressure loss for the introduced into the combustion chamber Combustion air is created.
  • this object is achieved by a along an axis-directed axial gas turbine comprising a compressor with a compressor exit region, an annular combustion chamber with a combustion air inlet area for From the compressor supplied combustion air, a combustion air supply line for the supply of combustion air from Compressor outlet area to the combustion air inlet area and with a turbine part, which has a turbine inlet area connected to the annular combustion chamber, wherein the combustion air supply line is formed so that a Verêtrluftdrall existing in the compressor outlet region the combustion air to the combustion air inlet area is essentially degraded and being on Combustion air inlet area around the whole circumference the ring combustion chamber extending swirl grille for the Issuing a twist on the combustion air arranged is.
  • the invention is based on the recognition that the use separate, single burners that run along the circumference open an annular combustion chamber, the flow velocity for in this burner entering combustion air through a maximum tolerable pressure drop is limited. Above one certain flow rate, the pressure drop is so great, that this negatively affects the efficiency of the gas turbine effect. However, a high flow rate is desirable as a result, on the one hand, flashbacks in the Burner can be avoided, on the other hand, an acoustic decoupling from the mixing process between fuel and air and the actual combustion is generated and therefore one Tendency to combustion vibrations is reduced.
  • the invention is an essential part of the pressure loss from the dissipation of the swirl in the combustion air caused.
  • fuel is uniform in the area of the swirl lattice distributed over the circumference of the swirl lattice in the Combustion air introduced.
  • Swirl blades of the swirl lattice at least partially as Fuel buckets formed over the fuel of the Combustion air can be supplied.
  • Such swirl vanes of the Twist lattices are then z. B. hollow, with her Inside fuel is supplied. Via openings on the surface the swirl vanes will then be the fuel in the Combustion air introduced. This leads to a homogeneous Distribution of the fuel in the combustion air.
  • Preferred dimensions include the distributable by the swirl lattice Twist of combustion air on a direction of rotation in the turbine inlet area in the direction of a diversion directed in the turbine inlet guide vanes is directed.
  • This smaller size is again advantageous in terms of a lower cooling air consumption and thus thus for efficiency as well as for lower nitrogen oxide emissions.
  • the gas turbine has a capacity greater than 50 Megawatts, with a flow rate generated at full load for the combustion air entering the swirl grid at a certain minimum speed.
  • Preferred dimensions is a ring around the annular combustion chamber circulating fuel supply line provided from the Fuel is supplied to the combustion air.
  • the fuel supply system be significantly simplified. This is about through the annular Fuel supply line possible from which the fuel z. B. is passed into the hollow swirl vanes.
  • Dependent from the pressure of the fuel at the respective position of the fuel Removal for a swirl bucket can use throttles for this Removal be provided so that, despite different Circumferential positions of the removal, and associated different Pressure in the fuel supply line, to the individual swirl blades supplied equal amounts of fuel can be.
  • a homogeneous distribution of Fuel is ensured over the circumference.
  • Preferred dimensions is the swirl lattice on the radially outer edge arranged the combustion air inlet region. hereby is at the radially inner edge of the annular combustion chamber, d. H. hub side, creates a single, annular rubström which effectively stabilizes combustion.
  • FIG. 1 shows a gas turbine 1.
  • the gas turbine 1 is along a turbine axis 3 directed. It includes a compressor 5, an annular combustion chamber 7 and a turbine part 9.
  • On one not shown shaft of the gas turbine 1 are wheel discs 11 for compressor blades 15 and wheel discs 13 arranged for turbine blades 17.
  • a swirl grid 33 is arranged, through which the combustion air 25 is introduced into the combustion chamber 7 becomes.
  • the combustion air 25 is in the swirl grid 33, as explained in more detail later with reference to FIG. 3, fuel admixed, the one of an annular fuel supply line 37th is taken, in turn, from a fuel main 35 is supplied.
  • the fuel air mixture passes as one Homogeneous premix in the annular combustion chamber 7, where there is a Hot gas 41 is burned.
  • a hub side annular Recirculation zone 43 stabilizes the combustion process.
  • the hot gas 41 then becomes a turbine inlet area 45 headed.
  • a first row of vanes 21 is arranged.
  • the Gas turbine 1 is enclosed by an outer housing 10.
  • the Ring combustion chamber 7 is with a ceramic heat shield brick lining 8 lined.
  • the swirl grating 33 Due to the arrangement of the swirl grating 33 is first at a such gas turbine ring combustion chamber 7 omitted it, to carry out the combustion process via separate burners, which are arranged along the circumference of the annular combustion chamber 7. Rather, it becomes a continuous entry for the combustion air 25 about that over the entire circumference of the annular combustion chamber 7 extending swirl grille 33 is provided. As a result, the hot gas 41 is an over the circumference of the annular combustion chamber 7 same spin granted, the only partially dissipates and to a considerable extent even in the turbine inlet area 45 is obtained. This leads to an im Comparison to the arrangement with separate burners, respectively have a single swirl lattice, to a lower one Pressure loss.
  • the hot gas 41 becomes along the circumference of the annular combustion chamber 7 evenly heated. Opposite an arrangement with separate burners, this has a uniform temperature distribution in the turbine inlet area result. The lesser Maximum temperatures thus take more agile Vanes 21, but also subsequent rows of blades. This means a reduced need for cooling air, that of the compressor air 25 removed and fed to the turbine blades becomes. This cooling air is no longer available to the combustion process available, resulting in reduced efficiency and also leads to higher nitrogen oxide emissions. The saving of cooling air by the temperature uniformity by means of the swirl lattice 33 thus has an increase in efficiency of the gas turbine 1 and reduced nitrogen oxide emissions result.
  • FIG. 2 shows a cross section through an annular combustion chamber 7.
  • the swirl lattice 33 is arranged in a circle.
  • the swirl lattice 33 is formed of swirl vanes 61, which will be described in more detail later with reference to FIG.
  • Around the swirl grille 33 extends an annular fuel supply line 37.
  • In this fuel supply line 37 is via a fuel main 35 from a fuel reservoir 53 fuel 54 initiated. From the fuel supply line 37 is then via swirl vane feeders 51, each of the swirl vanes 61 is supplied with fuel.
  • pilot burners 57 are diffusion burners, d. H.
  • Fuel and combustion air are only in the Combustion zone mixed. This burn is richer Fuel and more stable than a lean premix combustion. Because with such a diffusion combustion more nitrogen oxide formation follows, their share is kept as low as possible.
  • the Pilot burners 57 serve primarily to stabilize the Premix combustion generated via the swirl lattice 33.
  • FIG. 3 shows a swirl blade 61.
  • the swirl blade 61 is hollow executed. Into the interior of the swirl blade 61 is Fuel 54 initiated. This fuel 54 exits Openings 63 on the airfoil surface of the swirl vane 61 and mixes with the combustion air 25. This becomes a radially homogeneous distribution of Fuel reaches 54.
  • By the introduction into a variety of swirl vanes 61 evenly over the circumference of the combustion chamber 7 is also distributed a homogeneous distribution of Fuel 54 and combustion air 25 in the circumferential direction reached.
  • FIG. 4 shows the flow of the gas turbine guide vane 21 in the turbine inlet region 45 from the hot gas 41.
  • the hot gas 41 has a twist 71. This corresponds to a decomposition of Flow direction of the hot gas 41 in an axial and a to vertical component, wherein a twisting reflector 72 is stretched is, which has in the direction of a Umlenksektors 73, by the deflection of the hot gas 41 through the vanes 21 is shown. Due to the sense of rotation of the swirl 71 Thus, the hot gas 41 is already a component in the direction the deflection by the vanes 21 issued. With that you can the vanes 21 are made smaller, which in turn reduces the cooling air requirement for the guide vane 21.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
EP03018412A 2003-08-13 2003-08-13 Turbine à gaz Withdrawn EP1507120A1 (fr)

Priority Applications (1)

Application Number Priority Date Filing Date Title
EP03018412A EP1507120A1 (fr) 2003-08-13 2003-08-13 Turbine à gaz

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP03018412A EP1507120A1 (fr) 2003-08-13 2003-08-13 Turbine à gaz

Publications (1)

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EP1507120A1 true EP1507120A1 (fr) 2005-02-16

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Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1816400A3 (fr) * 2006-02-02 2012-07-04 Rolls-Royce Deutschland Ltd & Co KG Chambre de combustion de turbine à gaz dotée d'une injection de carburant sur la totalité de l'anneau de chambre de combustion
JP2017053618A (ja) * 2015-09-09 2017-03-16 ゼネラル・エレクトリック・カンパニイ 環状の流路構成体を有するシステムおよび方法

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3299632A (en) * 1964-05-08 1967-01-24 Rolls Royce Combustion chamber for a gas turbine engine
US4455840A (en) * 1981-03-04 1984-06-26 Bbc Brown, Boveri & Company, Limited Ring combustion chamber with ring burner for gas turbines
US5791148A (en) * 1995-06-07 1998-08-11 General Electric Company Liner of a gas turbine engine combustor having trapped vortex cavity
US5839283A (en) * 1995-12-29 1998-11-24 Abb Research Ltd. Mixing ducts for a gas-turbine annular combustion chamber

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3299632A (en) * 1964-05-08 1967-01-24 Rolls Royce Combustion chamber for a gas turbine engine
US4455840A (en) * 1981-03-04 1984-06-26 Bbc Brown, Boveri & Company, Limited Ring combustion chamber with ring burner for gas turbines
US5791148A (en) * 1995-06-07 1998-08-11 General Electric Company Liner of a gas turbine engine combustor having trapped vortex cavity
US5839283A (en) * 1995-12-29 1998-11-24 Abb Research Ltd. Mixing ducts for a gas-turbine annular combustion chamber

Cited By (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1816400A3 (fr) * 2006-02-02 2012-07-04 Rolls-Royce Deutschland Ltd & Co KG Chambre de combustion de turbine à gaz dotée d'une injection de carburant sur la totalité de l'anneau de chambre de combustion
JP2017053618A (ja) * 2015-09-09 2017-03-16 ゼネラル・エレクトリック・カンパニイ 環状の流路構成体を有するシステムおよび方法
CN106524224A (zh) * 2015-09-09 2017-03-22 通用电气公司 具有环形流动路径架构的系统和方法
EP3150917A3 (fr) * 2015-09-09 2017-07-12 General Electric Company Système et procédé de combustion ayant une architecture de trajet d'écoulement annulaire
US10465907B2 (en) 2015-09-09 2019-11-05 General Electric Company System and method having annular flow path architecture
CN106524224B (zh) * 2015-09-09 2020-07-14 通用电气公司 具有环形流动路径架构的系统和方法

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