WO2013050105A1 - Chambre de combustion pour une turbine à gaz - Google Patents

Chambre de combustion pour une turbine à gaz Download PDF

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Publication number
WO2013050105A1
WO2013050105A1 PCT/EP2012/003867 EP2012003867W WO2013050105A1 WO 2013050105 A1 WO2013050105 A1 WO 2013050105A1 EP 2012003867 W EP2012003867 W EP 2012003867W WO 2013050105 A1 WO2013050105 A1 WO 2013050105A1
Authority
WO
WIPO (PCT)
Prior art keywords
combustion chamber
inlet opening
inlet openings
extension direction
inlet
Prior art date
Application number
PCT/EP2012/003867
Other languages
German (de)
English (en)
Inventor
Stefan Czerner
Stefan KUNTZAGK
Original Assignee
Lufthansa Technik Ag
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 Lufthansa Technik Ag filed Critical Lufthansa Technik Ag
Publication of WO2013050105A1 publication Critical patent/WO2013050105A1/fr

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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
    • F23R3/04—Air inlet arrangements
    • F23R3/06—Arrangement of apertures along the flame tube
    • 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
    • F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
    • F23R2900/00019—Repairing or maintaining combustion chamber liners or subparts
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • the invention relates to a combustion chamber for a gas turbine with the features of the preamble of claim 1. Furthermore, the invention relates to a corresponding method for combustion chamber repair or production with the features of the preamble of claim 12.
  • combustion chamber walls are not sufficiently cooled, or if the flame geometry is unfavorable, burnings of the combustion chamber walls, in particular the popping of ceramic protective layers and the melting or increased oxidation or sulphiding of the combustion chamber wall can occur. Such unwanted burns can cause premature removal and repair of the combustion chamber, which is associated with high costs.
  • thermodynamic efficiency can be reduced by such burns, whereby more fuel is consumed in the rule.
  • Such unwanted combustion of the combustion chamber walls can be caused by the fact that the air cooling used for cooling the combustion chamber is canceled locally or
  • FIG. 1 shows a combustion chamber 6 of a ner gas turbine.
  • the outer wall 7 and the inner wall 8 surround the combustion chamber interior 5.
  • wall flows 1 are provided. Transverse to these wall flows 1, high pressure air is introduced into the combustion chamber 6 via inlet ports 4 to provide the oxygen for combustion.
  • the air of this inflowing air column 2 is also referred to as combustion air.
  • the flow conditions at an inlet opening 4 are shown by way of example in FIG. 8 in another perspective. It can be seen how the wall flows 1 a circular inlet opening 4, through which the air column 2 is blown into the combustion chamber interior 5, flow around. Downstream, behind the inlet opening 4, turbulences 3 are formed.
  • the damage mechanism can therefore be summarized in such a way that due to the turbulence after the inlet openings of the combustion air of the fuel or the flame enters more heat in the wall, as in other areas.
  • the materials fail, cracks occur and / or oxidation and / or sulfidation occurs with subsequent detachment of material.
  • the service life / function of the combustion chamber is thereby degraded in such a way that, for example, aircraft engines must be removed as a result of these deteriorations. Due to the reduced service life and / or reduced efficiencies of the turbomachine, it sometimes comes to considerable economic losses.
  • the invention is therefore based on the object to provide a combustion chamber in which there are improved flow conditions. Furthermore, the invention has for its object to provide a corresponding method for combustion chamber repair or manufacture.
  • a combustion chamber for a gas turbine with a combustion chamber interior is ignited in the supplied fuel, and an outer wall and an inner wall, which surround the combustion chamber, proposed according to the invention, wherein the outer and / or inner wall inlet openings, via which Combustion oxygen is supplied and wherein the geometry of at least one inlet opening is arranged so that it extends in at least one spatial direction farther than in a spatial direction perpendicular thereto.
  • the spatial direction in which the inlet opening extends further is referred to below as the main extension direction L.
  • the space direction perpendicular thereto is referred to as the lateral extension direction B.
  • the main extension direction L is the spatial direction in which the inlet opening has the greatest extent in which it is the longest. This is illustrated in FIG. 5, which will be described later.
  • the spatial directions of the inlet openings are to be understood as meaning the directions which lie in the plane perpendicular to the flow thread of the combustion air flowing through the inlet opening. Accordingly, the spatial directions refer to the cross-section of the combustion air and its geometry.
  • the length of the main extension direction L of at least one inlet opening is greater than 5 mm, preferably greater than 10 mm.
  • the length of the Mauerstre- ckungscardi B at least one inlet opening is greater than 3 mm, preferably greater than 6 mm.
  • the size of the inlet openings is important because the problems described above are more pronounced for larger inlet openings due to the greater turbulence. Thus, an improvement in the flow conditions at larger inlet openings has an increasingly positive effect.
  • Inlet openings in the sense of the invention are only those inlet openings which supply the air or the oxygen for the
  • cooling air openings which are mainly used for cooling of combustion chamber components, for example for the production of wall currents or film cooling, and at best afford only a negligible proportion of the oxygen supply, are not inlet openings in the context of this invention.
  • a major advantage of the combustion chamber according to the invention is that the course of the wall flows can be influenced in such an advantageous manner that it can not lead to the described defective burns of the combustion chamber walls or that such burns can be reduced. This is achieved by improving the flow conditions, in particular the wall flow conditions in the region of inlet openings.
  • the geometry of the inflowing air column is advantageously designed directly, since this naturally depends on the geometry of the inlet opening.
  • such geometries are shown for example in the figure 6, which will be described later.
  • Preference is given to rhombic, triangular, kite and drop-shaped geometries as well as asymmetrical or irregular geometries.
  • Further preferred are geometries in which the edges are rounded. Inlet openings with such geometries have increased resistance to cracking. The rounded edges further lead to a fanning of the introduced flow, which in individual cases can have a positive influence on the formation of flames.
  • At least one inlet opening has a ratio of main extension direction L to secondary extension direction B, which lies in a range between 1.2: 1 to 40: 1.
  • a range between 1.4: 1 to 20: 1 is preferred, more preferably a range between 1.6: 1 to 10: 1.
  • At least one inlet opening is elliptical.
  • elliptical are understood to mean geometries that resemble an ellipse or represent an ellipse, but no circles or substantially circular geometries.
  • Ellipse-shaped can also include, for example, olive-shaped geometries. In an olive-shaped geometry, the areas of two opposite vertices are significantly flattened and less round. As a result, the olive shape differs somewhat from the ideal elliptical geometry, but it is still elliptical. It is advantageous if the geometry of the at least one inlet opening changes in the direction of its main extension direction L. By elliptical inlet openings or inlet openings, which have an elliptical geometry, the flow conditions can be further influenced advantageous.
  • the main extension direction L of at least one inlet opening runs essentially parallel to the fluid axis.
  • These inlet openings are referred to below as Type A.
  • the mixing of the combustion air and fuel during combustion can be increased by the combustion chamber according to the invention.
  • the main extension direction L 'of at least one inlet opening runs essentially transversely to the fluid axis.
  • Such inlet openings are referred to below as Type B.
  • Inlet openings whose main extension direction is transverse to the fluid axis running - Type B - lead to increased turbulence.
  • such turbulence can have a negative effect on the components in some places, for example causing burns of the component walls.
  • at certain positions - preferably those less susceptible to burns - it may be advantageous to increase the strength of the turbulences in a targeted manner.
  • Such a reinforcement of the turbulence can for improved mixing of the combustion air with the fuel used or
  • inlet openings whose main extension direction L extends substantially parallel to the fluid axis further inlet openings are provided whose main extension direction L 'to the main extension direction L of the first inlet openings at an angle between 60 to 120 °, preferably about 90 ° .
  • specially adapted flow conditions can be set. It is advantageous if the respectively selected inlet openings are distributed with specific geometries over the entire circumference equidistant and for each annularly divided sections of the combustion chamber each have a kind of opening is provided.
  • a section with type B inlet openings adjoins a section with type A inlet openings. Further preferably, this is followed by a section with type A inlet openings.
  • inflowing air is providable by the passage of at least one inlet opening with a swirl.
  • This is preferably achieved by changing the cross-section of at least one inlet opening over the course of the opening depth, preferably the cross-section passes through a rotation.
  • the inlet opening on the air outflow side is rotated in relation to the air inflow side.
  • the type C of inlet openings This represents a further characteristic, the type C of inlet openings.
  • the exiting air column here has a twist and swirls strongly in the combustion chamber interior, as a result of which the fuel is mixed with the air to an above-average degree, which promotes combustion.
  • the difference to type B lies in the fact that the type C only produces turbulences in the middle of the combustion chamber, to the
  • Combustor walls type C - like the type A - leads to a reduced wake turbulence of the wall currents.
  • Inlet ports of the type C combine the advantageous vortex reduction in the wall area - the type A - with the improved combustion by increased turbulence in the combustion chamber flame - the type B -, with the additional swirl the targeted turbulence and thus the
  • Combustion can be further improved.
  • the swirl generated by the type C additionally generates turbulence in the center of the combustion chamber.
  • the flame can protrude to the end of the combustion chamber or close to the high-pressure turbine and damage the components there.
  • a section with type C inlet openings is arranged in front of and / or behind a section of type A inlet openings. It is preferably arranged between two sections of type A.
  • the overall efficiency of the gas turbine can be increased, thereby reducing emissions and saving fuel.
  • a method for combustor repair or production is also proposed, wherein in the course of repair or manufacture inlet openings of a combustion chamber are designed such that they have a geometry with the features described.
  • a combustor can be fitted directly to the specially shaped inlet ports, such as through bores, laser cutting, or as part of a casting process.
  • combustors having only circular inlet ports may be subsequently modified to have the advanced inlet ports. For example, holes can be widened to obtain elliptical inlet openings from circular inlet openings. By cutting processes, the geometries of the inlet openings can also be changed accordingly.
  • FIG. 1 is a schematic representation of a combustion chamber according to the prior art
  • Fig. 2 is a simplified representation of an inventive
  • FIG. 3 is a partial view of a combustion chamber outer wall with inlet openings according to the invention
  • Type B with the geometry of a rounded diamond
  • FIG. 1 shows the combustion chamber 6 of a gas turbine. This may be a stationary or a gas turbine provided for an aircraft. In this embodiment, it is a typical combustion chamber 6 from an aircraft engine.
  • the combustion chamber 6 has a combustion chamber interior 5, which is surrounded by an inner wall 8 and an outer wall 7. Between the two walls at the level of the fluid axis 12 is a fuel feed 9, through which the fuel required for combustion is admitted.
  • the outer and inner walls 7,8 have circular inlet openings 4, via which the combustion under high pressure (about 30 to 40 bar) air is supplied. Wall flows 1 are provided transversely to this incoming combustion air in order to cool the combustion chamber walls.
  • Such a combustion chamber with circular inlet openings is described for example in the document GB 2 465 853 A.
  • FIG. 8 shows a circular inlet opening 4, through which combustion air flows as air column 2 into the combustion chamber interior 5 and which is flowed around by the wall flows 1 described. Downstream downstream of the air column 2 or the inlet opening 4 - caused by the flow around the inlet opening or the air column 2 - 3 Verwirbe- ments from which are associated with the described losses of cooling capacity and the resulting harmful consequences.
  • the combustion chamber 6 shown in FIG. To illustrate the geometries and orientations of the inlet openings 4, the representation of the combustion chamber 6 with the inlet openings 4 is simplified and not true to scale. Combustion air 11 flows via the inlet openings 4 into the combustion chamber 6.
  • the wall angles 10 of the inlet openings 4 can be adjusted so that the combustion air 11 flows into the combustion chamber 6 at advantageous angles.
  • the geometry of the inlet openings is not circular, but various non-circular inlet openings are provided.
  • the Inlet openings 4 a drop-shaped 43 geometry.
  • the inlet openings 4 have an a-symmetrical geometry. Accordingly, type A inlet openings are provided in both sections.
  • the inlet openings 4 have an elliptical 40 geometry.
  • the peculiarity of the second section is that it is a combustion chamber section with inlet openings of the type C, the geometry or the orientation of the inlet openings 4 is therefore not constant over the depth of the inlet opening 4.
  • the alignment is-substantially similar to the fluid axis 12 -as in the first and third sections.
  • the geometry of the same inlet openings rotated by about 90 °, with the result that the combustion air is twisted as it flows through these inlet openings 4.
  • FIG. Such an outside of a combustion chamber 6 is shown in FIG. In this example, it is the outer wall of the combustion chamber 6. There are three sections shown, which are arranged one behind the other. In the first section, drop-shaped 43 inlet openings 4 are provided, in the second section elliptical 40 inlet openings 4 and in the third section the inlet openings 4 have the geometry of a rounded rhombus 48a.
  • FIGS. 4 and 5 both show the advantageous effect of inlet openings 4, in which the main extension direction L runs essentially parallel to the fluid axis 12. Due to the fact that individual segments of the fuel overlap karmmerwandung and the outer and inner walls slightly tapered overall in the flow direction in some combustion chambers, certain angular differences between the main extension direction L and fluid axis 12 may occur so that they do not exactly parallel, but only substantially parallel to each other. In an extension of their directions by imaginary lines, would be a meeting place of the imaginary lines outside the combustion chamber interior. 5
  • the lateral extension direction B is thus preferably transverse to the fluid axis 12. Shown is an elliptical 40 inlet opening 4 in FIG. 4 and an asymmetrical inlet opening 4 in FIG. 5.
  • the wall flow 1 flows around the inlet opening 4 or the air column flowing through the inlet opening 4 2, without causing turbulence 3 occur.
  • the discharge air 3a is less turbulent and thus the cooling effect of the wall flow 1 is not reduced.
  • FIG. 6 shows further advantageous geometries.
  • the inlet openings 4 may, for example, be diamond-shaped 48, triangular 41, kite-shaped 42, potato-shaped 44, olive-shaped 40a or drop-shaped 43.
  • house-shaped 46, arrow-shaped 47 or asymmetrical 45 geometries may be advantageous.
  • rounded diamonds 48a or rounded triangles 41a may be provided as the inlet opening geometry. All geometries have in common that they extend at least in one spatial direction further than in a direction perpendicular spatial direction.
  • the ratio of main extension direction L to secondary extension direction B is here preferably in a range between 1.2: 1 to 40: 1.
  • FIG. 7 shows an inlet opening 4 with the geometry of a rounded diamond 48a.
  • the main extension direction L ' is substantially transverse to the fluid axis 12. This has the consequence that the inflowing air behind the inlet port 4 amplified turbulence 3 has.
  • these turbulences 3 can increase the mixing of combustion air 11 and fuel. However, this is usually not advantageous for areas close to the combustion chamber walls, but for areas that are close to the center of the combustion flame.
  • a section with inlet openings 4 of the type C is preferably provided, in which at the same time the turbulences 3 can be reduced in the vicinity of the wall and increased in the central combustion chamber interior.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)

Abstract

Chambre de combustion (6) pour une turbine à gaz, qui comporte un espace interne (5) dans lequel du carburant introduit est enflammé, ainsi qu'une paroi externe (7) et une paroi interne (8) qui entourent l'espace interne (5) de la chambre de combustion, la paroi externe et/ou interne (7, 8) comportant des orifices d'entrée (4) par lesquels de l'oxygène est introduit pour la combustion. Selon l'invention, la forme d'au moins un orifice d'entrée (4) étant ainsi conçue qu'elle s'étend plus loin dans au moins une direction de l'espace que dans une direction de l'espace perpendiculaire à ladite direction.
PCT/EP2012/003867 2011-10-06 2012-09-17 Chambre de combustion pour une turbine à gaz WO2013050105A1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
DE102011114928.0 2011-10-06
DE102011114928A DE102011114928A1 (de) 2011-10-06 2011-10-06 Brennkammer für eine Gasturbine

Publications (1)

Publication Number Publication Date
WO2013050105A1 true WO2013050105A1 (fr) 2013-04-11

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WO (1) WO2013050105A1 (fr)

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3087266A4 (fr) * 2013-12-23 2017-05-17 United Technologies Corporation Configuration de trous de dilution à écoulements multiples pour un moteur à turbine à gaz
US10670267B2 (en) 2015-08-14 2020-06-02 Raytheon Technologies Corporation Combustor hole arrangement for gas turbine engine
US11774100B2 (en) 2022-01-14 2023-10-03 General Electric Company Combustor fuel nozzle assembly

Families Citing this family (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR3098569B1 (fr) * 2019-07-10 2021-07-16 Safran Aircraft Engines Paroi annulaire pour chambre de combustion de turbomachine comprenant des trous primaires, des trous de dilution et des orifices de refroidissement inclines

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DE1800612A1 (de) * 1968-10-02 1970-05-27 Hertel Dr Ing Heinrich Brennkammer,insbesondere fuer Strahltriebwerke
US20060130486A1 (en) * 2004-12-17 2006-06-22 Danis Allen M Method and apparatus for assembling gas turbine engine combustors
EP1840466A1 (fr) * 2006-03-30 2007-10-03 Snecma Configuration d'ouvertures de dilution dans une paroi de chambre de combustion de turbomachine
US20100031664A1 (en) * 2006-12-22 2010-02-11 Edward John Emilianowicz Combustor liner replacement panels
GB2465853A (en) 2008-12-08 2010-06-09 Gm Global Tech Operations Inc A double clutch transmission
US20100242483A1 (en) * 2009-03-30 2010-09-30 United Technologies Corporation Combustor for gas turbine engine
WO2011015543A1 (fr) * 2009-08-04 2011-02-10 Snecma Chambre de combustion de turbomachine comprenant des orifices d'entree d'air ameliores
US20110048024A1 (en) * 2009-08-31 2011-03-03 United Technologies Corporation Gas turbine combustor with quench wake control

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DE102006026969A1 (de) * 2006-06-09 2007-12-13 Rolls-Royce Deutschland Ltd & Co Kg Gasturbinenbrennkammerwand für eine mager-brennende Gasturbinenbrennkammer
DE102007018061A1 (de) * 2007-04-17 2008-10-23 Rolls-Royce Deutschland Ltd & Co Kg Gasturbinenbrennkammerwand
US7984615B2 (en) * 2007-06-27 2011-07-26 Honeywell International Inc. Combustors for use in turbine engine assemblies

Patent Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE1800612A1 (de) * 1968-10-02 1970-05-27 Hertel Dr Ing Heinrich Brennkammer,insbesondere fuer Strahltriebwerke
US20060130486A1 (en) * 2004-12-17 2006-06-22 Danis Allen M Method and apparatus for assembling gas turbine engine combustors
EP1840466A1 (fr) * 2006-03-30 2007-10-03 Snecma Configuration d'ouvertures de dilution dans une paroi de chambre de combustion de turbomachine
US20100031664A1 (en) * 2006-12-22 2010-02-11 Edward John Emilianowicz Combustor liner replacement panels
GB2465853A (en) 2008-12-08 2010-06-09 Gm Global Tech Operations Inc A double clutch transmission
US20100242483A1 (en) * 2009-03-30 2010-09-30 United Technologies Corporation Combustor for gas turbine engine
WO2011015543A1 (fr) * 2009-08-04 2011-02-10 Snecma Chambre de combustion de turbomachine comprenant des orifices d'entree d'air ameliores
US20110048024A1 (en) * 2009-08-31 2011-03-03 United Technologies Corporation Gas turbine combustor with quench wake control

Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3087266A4 (fr) * 2013-12-23 2017-05-17 United Technologies Corporation Configuration de trous de dilution à écoulements multiples pour un moteur à turbine à gaz
US10386070B2 (en) 2013-12-23 2019-08-20 United Technologies Corporation Multi-streamed dilution hole configuration for a gas turbine engine
US10670267B2 (en) 2015-08-14 2020-06-02 Raytheon Technologies Corporation Combustor hole arrangement for gas turbine engine
EP3130855B1 (fr) * 2015-08-14 2022-06-15 Raytheon Technologies Corporation Paroi de chambre de combustion de turbine à gaz comprenant une agencement de perforations
US11774100B2 (en) 2022-01-14 2023-10-03 General Electric Company Combustor fuel nozzle assembly

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