EP1656497B1 - Diffuseur situe entre le compresseur et la chambre de combustion d'une turbine a gaz - Google Patents

Diffuseur situe entre le compresseur et la chambre de combustion d'une turbine a gaz Download PDF

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Publication number
EP1656497B1
EP1656497B1 EP04741084A EP04741084A EP1656497B1 EP 1656497 B1 EP1656497 B1 EP 1656497B1 EP 04741084 A EP04741084 A EP 04741084A EP 04741084 A EP04741084 A EP 04741084A EP 1656497 B1 EP1656497 B1 EP 1656497B1
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EP
European Patent Office
Prior art keywords
combustion chamber
turbine
diffuser
longitudinal axis
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.)
Not-in-force
Application number
EP04741084A
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German (de)
English (en)
Other versions
EP1656497A1 (fr
Inventor
Peter Tiemann
Reinhard MÖNIG
Christian Cornelius
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Siemens AG
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Siemens AG
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Publication date
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Priority to PL04741084T priority Critical patent/PL1656497T3/pl
Priority to EP04741084A priority patent/EP1656497B1/fr
Publication of EP1656497A1 publication Critical patent/EP1656497A1/fr
Application granted granted Critical
Publication of EP1656497B1 publication Critical patent/EP1656497B1/fr
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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/08—Cooling; Heating; Heat-insulation
    • F01D25/12—Cooling
    • 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
    • F23R3/005—Combined with pressure or heat exchangers

Definitions

  • the invention relates to a gas turbine with an annular combustion chamber and one of these upstream, substantially parallel to a turbine longitudinal axis and flowed from this less than the annular combustion chamber spaced diffuser, in which a compressed gas at a branch point in partial flows can be divided.
  • Gas turbines are used in many areas to drive generators or work machines.
  • the energy content of a fuel is used to generate a rotational movement of a turbine shaft.
  • the fuel is burned in a combustion chamber, compressed air being supplied by an air compressor.
  • the working medium produced in the combustion chamber by the combustion of the fuel, under high pressure and at high temperature, is guided via a turbine unit arranged downstream of the combustion chamber, where it relaxes to perform work.
  • Cooling of the affected components in particular of running and / or vanes of the turbine unit, provided. Furthermore, it can be provided to cool the combustion chamber with a coolant, in particular cooling air.
  • a gas turbine which has a combustion chamber upstream and opening into a diffuser air compressor.
  • a partial flow of the compressed air can be branched out of the diffuser and used for cooling structural parts, for example turbine blades of the gas turbine.
  • thede povertya stoodeist from the diffuser is only suitable for a branch of a relatively small partial flow from the air flow leaving the air compressor.
  • the main air flow conducted through the diffuser is deflected in the diffuser in the direction of the combustion chamber and supplied to it as combustion air.
  • a cooling of the downstream of the diffuser that is, based on the flow direction of the working medium flowing through the turbine downstream components is thus limited possible.
  • DE 196 39 623 discloses a gas turbine with a diffuser, in which the removal of the cooling air takes place by means of a tube projecting into the outlet of the diffuser.
  • the compressed air used for combustion in an annular combustion chamber is thereby diverted by means of a C-shaped plate in the direction of the burner. Both when removing the cooling air as well as in the leadership of the burner air flow losses can occur, which should be avoided.
  • the invention has for its object to provide a equipped with an annular combustor compact gas turbine, which allows a favorable flow guidance of the compressor air for a particularly uniform and effective cooling of thermally loaded components.
  • a gas turbine with the features of claim 1.
  • the gas turbine on an annular combustion chamber and an upstream of this annular diffuser, which is at least partially disposed between the turbine longitudinal axis and the annular combustion chamber.
  • the diffuser which can be flowed essentially parallel to the turbine longitudinal axis, a compressed gas can be divided into a plurality of partial flows.
  • the diffuser has a main deflection region, which is directed at an acute angle from the turbine longitudinal axis in a pioneering manner onto the inner wall of the annular combustion chamber.
  • the main deflection region is followed in the direction of the gas flowing through the diffuser, in particular air, a branching point, at which the gas flowing through the diffuser can be divided into partial flows by means of a flow dividing element.
  • the annular and in cross-section wedge-shaped flow dividing element is arranged between the two diverging walls of the diffuser - the radially inner inner wall and the radially outer wall lying outside.
  • Two deflecting flanks opposite the walls of the diffuser converge towards each other at an acute angle and meet at the branching point. There they include an angle bisector, which intersects the turbine longitudinal axis at an acute pitch angle greater than 15 °.
  • the Hauptablenk Scheme is seen in the axial direction behind the compressor and in front of the annular combustion chamber, whereas the flow dividing element between the annular combustion chamber and the turbine longitudinal axis is arranged.
  • This geometry allows for the gas turbine a compact and in particular a shortened in the axial direction design. Furthermore, the flow losses in the compressed refrigerant partial streams are reduced.
  • the two partial streams divided in the diffuser are also used in connection for combustion.
  • the outer wall of the diffuser and the outer deflecting flank of the flow-dividing element lying opposite to it extend approximately perpendicular to the turbine longitudinal axis behind the branching point. This ensures a low-loss supply of the outer partial flow to the outer flow passage space. A short and direct supply of the partial flow is achieved accordingly.
  • the supply of the outer combustion chamber shell is quite simple.
  • the individual flute-shaped combustion chambers are spaced apart on a ring concentrically enclosing the turbine longitudinal axis in the circumferential direction. The supply of cooling air to the radially outer combustion chamber shells can then take place between the individual Can combustion chambers.
  • a low-loss supply of the inner partial flow to the inner flow passage space is ensured by the inner wall of the diffuser and the opposite inner deflecting edge of the flow dividing element extends approximately parallel to the turbine longitudinal axis.
  • a wavy guide is proposed for the inner partial flow, which achieves an improvement over a linear guide in comparison to a straight guide with regard to the pressure losses and the flow losses in the partial flow.
  • the compressed gas which leaves the diffuser at this point, is conducted directly into a flow transfer space at the branch point, which passes the fluidic connection to the wall cooling space produces the annular combustion chamber.
  • the flow transfer space preferably adjoins the outside of the combustion chamber wall, so that an additional cooling of the combustion chamber wall is achieved as a result.
  • the ring combustion chamber is preferably formed closed coolable.
  • combustion air is preferably performed as a cooling medium in countercurrent to the flue gas through a wall space of the annular combustion chamber.
  • the combustion air flowing through the combustion chamber wall is preferably identical here, at least with a partial flow of the compressed air, which has previously flowed through the diffuser.
  • the air flowing through the diffuser is supplied completely to the wall of the annular combustion chamber as cooling air and further to the annular combustion chamber as combustion air.
  • the division of the air flow at the branch point of the diffuser serves to provide several parts of the annular combustion chamber, such as an inner shell and an outer shell, evenly with cooling air.
  • the wall angle of the annular combustion chamber is understood to mean that angle which the combustion chamber rear wall encloses with the turbine longitudinal axis.
  • a particularly uniform all-round cooling of the combustion chamber wall is preferably achieved in that the pitch angle of the flow dividing element deviates from the wall angle of the combustion chamber rear wall by not more than 20 °, in particular by not more than 15 °.
  • a pipe communicating with the lower part of the channel is provided for the removal of cooling air for the turbine.
  • This allows a further division of the compressor air flow. If the tube protrudes into the lower part of the channel and faces with its pipe opening to the flow, the extraction of turbine cooling air is particularly favorable.
  • the advantage of the invention lies in the fact that in a gas turbine compressed air, which serves as a cooling and then combustion air is supplied with low pressure loss of an air compressor through a compact diffuser of the annular combustion chamber, wherein a flow divider at the outlet of the diffuser uniform cooling air to the annular combustion chamber causes.
  • the gas turbine 1 has a compressor 2 for combustion air, an annular combustion chamber 4 and a turbine 6 for driving the compressor 2 and a generator or a working machine (not shown).
  • the turbine 6 and the compressor 2 are arranged on a common, also called turbine rotor turbine shaft 8, with which the generator or the working machine is connected, and which is rotatably mounted about its central axis 9.
  • the annular combustion chamber 4 is equipped with a number of burners 10 for the combustion of a liquid or gaseous fuel. It is also provided on its combustion chamber wall 23 with a wall lining 24.
  • the turbine 6 has a number of rotatable blades 12 connected to the turbine shaft 8.
  • the blades 12 are arranged in a ring on the turbine shaft 8 and thus form a number of blade rows.
  • the turbine 6 comprises a number of fixed vanes 14, which are also secured in a ring shape with the formation of vane rows on an inner housing 16 of the turbine 6.
  • the blades 12 serve to drive the turbine shaft 8 by momentum transfer from the turbine 6 flowing through the flue gas or working medium M.
  • the vanes 14, however, serve to guide the flow of the working medium M between two seen in the flow direction of the working medium M consecutive blade rows or blade rings.
  • a successive pair of a ring of vanes 14 or a row of vanes and a ring of blades 12 or a blade row is also referred to as a turbine stage.
  • Each vane 14 has a platform 18, also referred to as a vane foot 19, which is intended to fix the respective vane 14 in the gas turbine 1.
  • Each blade 12 is attached to the turbine shaft 8 in an analogous manner via a blade root 19, also referred to as a platform 18, the blade root 19 each carrying a profiled blade 20 extended along a blade axis.
  • each guide ring 21 on the inner housing 16 of the turbine 6 is arranged between the spaced-apart platforms 18 of the guide vanes 14 of two adjacent rows of guide vanes.
  • the outer surface of each guide ring 21 is also exposed to the hot, the turbine 6 flowing through the working medium M and spaced in the radial direction from the outer end 22 of the blade 12 opposite him through a gap.
  • the guide rings 21 arranged between adjacent rows of guide blades serve in particular as cover elements which protect the inner wall 16 or other housing installation parts from thermal overload by the hot working medium M flowing through the turbine 6.
  • the gas turbine 1 is designed for a comparatively high outlet temperature of the working medium M emerging from the annular combustion chamber 4 from about 1200 ° C. to 1300 ° C.
  • the combustion chamber wall 23 can be cooled with cooling air compressed in the compressor 2 as coolant K. Between the combustion chamber wall 23 and the wall lining 24, cooling air K flows in a wall space or wall lining space 26 in countercurrent to the working medium M onto the burner 10.
  • the cooling air K which also serves as combustion air, is passed from the compressor 2 through a diffuser 27 in the direction of the annular combustion chamber 4. By the diffuser 27, the cooling and combustion air K defined split on the one hand to an outer combustion chamber shell 28 and on the other hand, an inner combustion chamber shell 29 is supplied.
  • the diffuser 27 has a main deflection region 30, which adjoins the compressor 2.
  • the compressed cooling air K flows out of the compressor 2 parallel to the central axis or turbine longitudinal axis 9 and into the main deflection region 30 of the diffuser 27.
  • the main deflecting region 30 of the diffuser 27 arranged axially between the compressor 2 and the annular combustion chamber 4 extends radially outwards under cross-sectional expansion, ie. away from the turbine longitudinal axis 9. As a result, the flow velocity of the compressed gas used as coolant K is reduced in the main deflection region 30. If there is a flow separation on the inner wall and outer wall of the diffuser 27, such a separation occurs only at low flow velocity and correspondingly low pressure loss.
  • a flow dividing element 32 is disposed adjacent to the outer combustion chamber shell 29.
  • the arranged between the annular combustion chamber 4 and the turbine longitudinal axis 9 flow dividing element 32 has an approximately triangular in cross-section, also referred to as a dividing fork 33 shape with an outer Ablenkflanke 34 and an inner Ablenkflanke 35.
  • the deflection flanks 34, 35 converge toward a division tip 36 directed towards the main deflection region 30 and enclose an acute angle of less than 90 °, in particular an angle of 60 °, in the division tip 36.
  • the dividing point or edge 36 forming a branching point divides the cooling air K flowing through the main deflecting region 30 of the diffuser 27 approximately equally into an outer cooling air flow K a and an inner cooling air flow K i .
  • the outer cooling air flow K a is fed through an outer flow transfer chamber 37 of an outer combustion chamber shell 28, while the inner cooling air flow K i is fed via an inner flow transfer chamber 38 of the inner combustion chamber shell 29.
  • the diffuser 27 dividing the cooling air K at the flow dividing element 32 is also referred to as a split diffuser.
  • the cooling air K flowing through the main deflecting region 30 is directed approximately C-shaped radially, relative to the turbine longitudinal axis 9, outwardly to the dividing point 36 of the flow dividing element 32.
  • a line extending as an angle bisector 39 between the curved Ablenkflanken 34,35 through the divisional peak 36 includes with the turbine longitudinal axis 9 a pitch angle ⁇ of about 45 °.
  • the bisector 39 includes an approximately right angle.
  • the inner cooling air flow K i is, starting from the division tip 36, forced by the inner Ablenkflanke 35 first in a horizontal flow direction, ie parallel to the turbine longitudinal axis 9 and further through the outside of the combustion chamber wall 23 radially inward, ie towards the turbine longitudinal axis 9, directed.
  • the inner cooling air flow K i is thus, initially still within the undivided in the main deflection region 30 cooling air K, guided in a roughly C-shaped curved path radially outward and thereby delayed and then guided in a direction in the opposite direction, approximately C-shaped curved path radially inward.
  • the flow through the diffuser 27 and further into the internal flow transfer space 38 describes approximately a double S-shaped path. The radii of curvature within this path are large enough to cause only small energy losses in the flow.
  • guide members or fixing members 41 are disposed both in the direction of the outer flow passage space 37 and the direction of the inner flow passage space 38.
  • the outer cooling air flow K a is guided by the dividing fork 33 radially, perpendicular to the turbine longitudinal axis 9, to the outside. In the course of the outer cooling air flow K a is guided past the outer combustion chamber shell 28 and introduced into the wall lining room or wall cooling space 26. Again, similar to the inner cooling air flow K i results in a flow guidance with large deflection radii, with no sudden cross-sectional enlargements occur. Due to the cooling air streams or partial streams K a , K i , the combustion chamber shells 28, 29 are also cooled from the outside.
  • the burner 10 is arranged approximately centrally in a combustion chamber rear wall 42.
  • a straight line passing through the combustion chamber rear wall 42 encloses the turbine longitudinal axis 9 with a wall angle ⁇ of approximately 45 °.
  • the wall angle ⁇ thus corresponds approximately to the pitch angle ⁇ .
  • the flow splitting element 32 arranged at an angle to the turbine longitudinal axis 9 at a pitch angle ⁇ splits the main deflecting region 30 into an upper sub-channel 43 and a lower sub-channel 44, which both have approximately the same cross-section.
  • offset arrangement of the flow dividing element 32 is also a targeted asymmetrical division of the cooling air flow in the diffuser 27 feasible, if, for example, the outer combustion chamber shell and the inner combustion chamber shell 29 have a different cooling air requirement.
  • the removal for turbine cooling air is effected by a projecting into the lower part of the duct 44 tube 45.
  • Whose end 46 is angled in the manner of a periscope and facing with its tube opening the inner air flow K i , so that a portion of the air flow K i can flow into the tube 45 as a turbine cooling air.
  • the turbine cooling air flows at the other end of the tube 45 into an annular channel 47 extending along the rotor, which leads the turbine cooling air to the turbine 6. There, it is used for cooling the rotor blades and guide vanes 12, 14.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

Abstract

L'invention concerne une turbine à gaz (1) comprenant une chambre de combustion (4) annulaire et un diffuseur (27) qui est implanté en amont de cette chambre, qui peut être parcouru de manière sensiblement parallèle à un axe longitudinale (9) de la turbine et qui est espacé de cet axe d'une distance au moins partiellement inférieure à celle de la chambre de combustion annulaire (4). Dans ce diffuseur, un gaz comprimé (K) est divisé en plusieurs flux partiels (Ki,Ka) au niveau d'un point de dérivation (36), au moins un flux partiel (Ki,Ka) étant un flux de gaz de refroidissement. Cette turbine à gaz présente dans le diffuseur (27) une zone de déviation principale (30) qui est inclinée de l'axe longitudinal (9) de la turbine à la chambre de combustion annulaire (4).

Claims (11)

  1. Turbine (1) à gaz ayant une chambre de combustion (4) annulaire inclinée par rapport à l'axe (9) longitudinal de la turbine,
    qui a une paroi (42) arrière de chambre de combustion, dans laquelle s'étend une ligne de paroi coupant l'axe (9) longitudinal de la turbine suivant un angle β aigu de paroi d'au moins 30°,
    comprenant un compresseur (2) en aval duquel, en technique des fluides, est monté dans la direction axiale un diffuseur (27) disposé au moins en partie entre la chambre de combustion (4) annulaire et l'axe (9) longitudinal de la turbine,
    caractérisée
    en ce que, dans le diffuseur (27), un gaz (K) comprimé peut être subdivisé en un point (36) de bifurcation par un élément (32) cunéiforme de division du courant formé par deux flancs (34, 35) de déviation en des sous-courants (Ki, Ka) ,
    dans lequel, au point (36) de bifurcation, les deux flancs (34, 35) de déviation font un angle plus petit que 90° et leur bissectrice coupe l'axe (9) longitudinal de la turbine suivant un angle α aigu supérieur à 15°, et
    dans lequel le diffuseur (27) a, en amont du point (36) de bifurcation, une zone (30) de déviation principale qui s'écarte suivant un angle aigu de l'axe (9) de la turbine longitudinal vers une coque (29) intérieure de la chambre de combustion (4) annulaire, s'étendant transversalement à la paroi (42) arrière de la chambre de combustion.
  2. Turbine (1) à gaz suivant la revendication 1,
    dans laquelle le flanc (34) de déviation extérieur délimitant le sous-courant (Ka) extérieur radialement et une paroi extérieure du diffuseur (27) opposée à ce flanc (34) de déviation s'étendent derrière le point (36) de bifurcation à peu près perpendiculairement à l'axe (9) longitudinal de la turbine.
  3. Turbine (1) à gaz suivant la revendication 1 ou 2,
    dans laquelle le flanc (35) de déviation intérieur délimitant le sous-courant (Ki) intérieur radialement et une paroi intérieure du diffuseur (27) opposée à ce flanc (35) de déviation s'étendent derrière le point (36) de bifurcation à peu près parallèlement à l'axe (9) longitudinal de la turbine.
  4. Turbine (1) à gaz suivant la revendication 3,
    dans laquelle le sous-courant (Ki) intérieur radialement peut passer après avoir quitter le diffuseur (27) de manière inclinée dans la direction de l'axe (9) longitudinal de la turbine.
  5. Turbine (1) à gaz suivant l'une des revendications 1 à 4,
    comprenant un espace (26) de refroidissement de paroi de la chambre de combustion (4) annulaire, constitué sous la forme d'une coquille (29) intérieure de chambre de combustion et d'une coquille 28 extérieure de chambre de combustion.
  6. Turbine (1) à gaz suivant la revendication 5
    comprenant un espace (37, 38) de passage du courant qui est voisin de la chambre de combustion (4) annulaire et qui met le diffuseur (27) en communication avec l'espace (26) de refroidissement de paroi.
  7. Turbine (1) à gaz suivant l'une des revendications 1 à 6,
    comprenant une chambre de combustion (4) annulaire refroidie en circuit fermé.
  8. Turbine (1) à gaz suivant l'une des revendications 1 à 7,
    dans laquelle la chambre de combustion (4) annulaire est refroidie par un procédé à contre-courant.
  9. Turbine (1) à gaz suivant l'une des revendications 1 à 8,
    dans laquelle l'angle α de la bissectrice diffère de pas plus de 20° de l'angle β de paroi.
  10. Turbine (1) à gaz suivant l'une des revendications 1 à 9,
    dans laquelle il est prévu pour prélever de l'air de refroidissement pour la turbine un tuyau (45) communiquant avec le sous-canal (44) inférieur.
  11. Turbine (1) à gaz suivant la revendication 10,
    dans lequel le tuyau (45) pénètre dans le sous-canal (44) inférieur et est tourné vers le courant par son ouverture de tuyau.
EP04741084A 2003-08-18 2004-07-16 Diffuseur situe entre le compresseur et la chambre de combustion d'une turbine a gaz Not-in-force EP1656497B1 (fr)

Priority Applications (2)

Application Number Priority Date Filing Date Title
PL04741084T PL1656497T3 (pl) 2003-08-18 2004-07-16 Dyfuzor umieszczony pomiędzy sprężarką i komorą spalania turbiny gazowej
EP04741084A EP1656497B1 (fr) 2003-08-18 2004-07-16 Diffuseur situe entre le compresseur et la chambre de combustion d'une turbine a gaz

Applications Claiming Priority (3)

Application Number Priority Date Filing Date Title
EP03018565A EP1508680A1 (fr) 2003-08-18 2003-08-18 Diffuseur situé entre le compresseur et la chambre de combustion d'une turbine à gaz
EP04741084A EP1656497B1 (fr) 2003-08-18 2004-07-16 Diffuseur situe entre le compresseur et la chambre de combustion d'une turbine a gaz
PCT/EP2004/007946 WO2005019621A1 (fr) 2003-08-18 2004-07-16 Diffuseur place entre le compresseur et la chambre de combustion d'une turbine a gaz

Publications (2)

Publication Number Publication Date
EP1656497A1 EP1656497A1 (fr) 2006-05-17
EP1656497B1 true EP1656497B1 (fr) 2006-11-02

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Application Number Title Priority Date Filing Date
EP03018565A Withdrawn EP1508680A1 (fr) 2003-08-18 2003-08-18 Diffuseur situé entre le compresseur et la chambre de combustion d'une turbine à gaz
EP04741084A Not-in-force EP1656497B1 (fr) 2003-08-18 2004-07-16 Diffuseur situe entre le compresseur et la chambre de combustion d'une turbine a gaz

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Application Number Title Priority Date Filing Date
EP03018565A Withdrawn EP1508680A1 (fr) 2003-08-18 2003-08-18 Diffuseur situé entre le compresseur et la chambre de combustion d'une turbine à gaz

Country Status (7)

Country Link
US (1) US8082738B2 (fr)
EP (2) EP1508680A1 (fr)
CN (1) CN100390387C (fr)
DE (1) DE502004001924D1 (fr)
ES (1) ES2275226T3 (fr)
PL (1) PL1656497T3 (fr)
WO (1) WO2005019621A1 (fr)

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EP1245804B1 (fr) * 2001-03-26 2006-05-24 Siemens Aktiengesellschaft Turbine à gaz
DE50107283D1 (de) * 2001-06-18 2005-10-06 Siemens Ag Gasturbine mit einem Verdichter für Luft
EP1400751A1 (fr) * 2002-09-17 2004-03-24 Siemens Aktiengesellschaft Chambre de combustion pour turbine à gaz
GB0229307D0 (en) * 2002-12-17 2003-01-22 Rolls Royce Plc A diffuser arrangement

Also Published As

Publication number Publication date
EP1508680A1 (fr) 2005-02-23
US8082738B2 (en) 2011-12-27
CN1836097A (zh) 2006-09-20
WO2005019621A1 (fr) 2005-03-03
ES2275226T3 (es) 2007-06-01
EP1656497A1 (fr) 2006-05-17
CN100390387C (zh) 2008-05-28
PL1656497T3 (pl) 2007-03-30
DE502004001924D1 (de) 2006-12-14
US20100257869A1 (en) 2010-10-14

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