EP2567071A1 - Zone de transition pour une chambre de combustion secondaire d'une turbine à gaz - Google Patents

Zone de transition pour une chambre de combustion secondaire d'une turbine à gaz

Info

Publication number
EP2567071A1
EP2567071A1 EP11716539A EP11716539A EP2567071A1 EP 2567071 A1 EP2567071 A1 EP 2567071A1 EP 11716539 A EP11716539 A EP 11716539A EP 11716539 A EP11716539 A EP 11716539A EP 2567071 A1 EP2567071 A1 EP 2567071A1
Authority
EP
European Patent Office
Prior art keywords
gas turbine
radially
cavity
wall
region
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Withdrawn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
EP11716539A
Other languages
German (de)
English (en)
Inventor
Thomas Heinz-Schwarzmaier
Marc Widmer
Selma Zahirovic
Paul Marlow
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.)
Ansaldo Energia IP UK Ltd
Original Assignee
Alstom Technology 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 Alstom Technology AG filed Critical Alstom Technology AG
Publication of EP2567071A1 publication Critical patent/EP2567071A1/fr
Withdrawn legal-status Critical Current

Links

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
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/023—Transition ducts between combustor cans and first stage of the turbine in gas-turbine engines; their cooling or sealings
    • 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
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • 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
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/005—Sealing means between non relatively rotating elements
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C6/00—Plural gas-turbine plants; Combinations of gas-turbine plants with other apparatus; Adaptations of gas-turbine plants for special use
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C6/00—Plural gas-turbine plants; Combinations of gas-turbine plants with other apparatus; Adaptations of gas-turbine plants for special use
    • F02C6/003—Gas-turbine plants with heaters between turbine stages
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/10—Stators
    • F05D2240/12—Fluid guiding means, e.g. vanes
    • F05D2240/127—Vortex generators, turbulators, or the like, for mixing

Definitions

  • the present invention relates to a gas turbine, in particular a special embodiment of the transition region between a secondary combustion chamber and a low-pressure turbine in a gas turbine.
  • Gas turbine may be configured with a single combustion chamber, but they may also have a so-called sequential combustion. In the latter, fuel is burned in a first combustion chamber and the combustion air is subsequently expanded via a first turbine, a high-pressure turbine. Behind the high-pressure turbine, the still hot combustion gases flow through a secondary combustion chamber in which fuel is again supplied and burned under auto-ignition. Behind this secondary combustion chamber, a low-pressure turbine is arranged, through which the combustion gases are expanded, optionally followed by a heat recovery system with steam generation.
  • the transition of the housing from combustion chamber to turbine is a critical area, because there are particularly complex temperature and pressure conditions in this.
  • the secondary combustion chamber which is normally formed as an annular combustion chamber, has a somewhat cupped outer boundary, an outer wall which consists of a heat-resistant material or is coated accordingly, and which is normally composed of individual segments.
  • the low pressure turbine in turn has a plurality of alternating rows of vanes and blades.
  • the first row of blades arranged immediately behind the secondary combustion chamber is typically a row of vanes with significant rotation of the blades with respect to the major axis direction.
  • the guide vanes are typically formed as a segment module, in which each vane on the inside an inner platform and on the outside has an outer platform, and these platforms then with their inner surface and the flow channel of the combustion air radially inside or radially outward limit.
  • this gap there is a gap between the inner wall segment of the secondary combustion chamber and the inner platform of the first blade row on the radially inner side of the annular flow channel and a gap on the radially outer side between the outer wall segment of the secondary combustion chamber and the outer platform of the first blade row.
  • this gap must have a certain width and can not be easily closed or completely bridged.
  • the problem with this gap, which forms a cavity which extends quite far radially outwards into further structural components of the housing, in particular on the radial outside, is the fact that it is also exposed to complex flow conditions, in particular in the region of one guide vane.
  • the present invention intervenes and goes a completely different way than the prior art.
  • no attempt should be made to close the gap as this results in the above-mentioned problems.
  • the gap is indeed a have certain width (in the axial direction), but it should be prevented by appropriate measures that hot air or combustion air can enter this gap easily and can trigger the above problems.
  • a gas turbine with a secondary combustion chamber and immediately downstream therefrom arranged first row of vanes of a low-pressure turbine wherein the radially outer boundary of the secondary combustion chamber is formed by at least one outer wall segment, which is attached to at least one radially outwardly disposed support member, wherein the Flow path of the hot gases in the region of the vane row is bounded radially on the outside by an outer platform which is at least indirectly attached to at least one vane carrier, and wherein between the wall segment and the outer platform a substantially radially extending gap-shaped cavity having a width B in the inlet region in the axial Direction is present in the range of 1-25 mm. Where the width B for the cold installation state is indicated. Depending on the size of the housing clearance and permitted tolerances, the width B is in a range of 2 - 15 mm.
  • this gap is not closed or extremely narrowed in terms of gap size, at least in the entry region, but instead, at least one step element is arranged in the inlet region, which width B in at least one, substantially perpendicular to the flow direction of the hot gas in the Cavity running stage reduced by at least 10%.
  • this step element which is arranged substantially immediately behind the actual inlet area (typically 10-50 mm radially outside the entrance gap), flow vortices are generated, which to a certain extent assume a sealing function and prevent the hot air can penetrate into the cavity in depth. Accordingly, it is also important that the stage can produce such swirls, and so preferably the stage is formed as a single stage which realizes the indicated reduction of at least 10% in one step. Typically, the step has substantially rectangular transition areas.
  • the step element is formed circumferentially with respect to the axis of the turbine. Accordingly, the step element is formed substantially as a circumferential, arranged in the cavity at one of its walls rib. It may be a single such step element arranged in the cavity, it can but also several such step elements are provided offset radially outwards. Accordingly, it is possible to expand the cavity again behind the first stage and to provide a second step element after this extension. This creates two vortices and ensures an increased sealing function. With sufficient width B of the cavity, at least one further step element can be arranged on the wall of the cavity opposite the first step element. Typically, the step elements face each other so that there is a constriction from both sides of the cavity.
  • the step element form circumferentially segmentally each blade is assigned radially outside such a segment (that is, lying between the segments areas of the cavity do not have a step element).
  • substantially all segments have a length in the direction of rotation relative to the circumferential distance p (pitch) of the guide vanes of 30-50% of the circumferential distance p.
  • each segmentally circumferentially formed step elements may, for example symmetrical to the vane (that is, circumferentially in the same mass from the radial position of the leading edge in a clockwise and counterclockwise extending) be assigned or arranged according to a radial offset of the bow wave, offset from the vanes.
  • a further preferred embodiment of the proposed gas turbine is characterized in that the step element is designed in the form of a rib which is placed on or integrally formed on the wall region of the outer platform bordering on the cavity and substantially rectangular in axial cross-section.
  • this has a length in the radial direction in the range of 10-100 mm, in particular preferably in the range of 20-50 mm.
  • this is preferably used in combination with a radially outwardly arranged in this wall region formed recess of an equal or greater length, whose radially outer side end is formed by a further step, so that radially Two or three vortices are created one behind the other and an increased sealing effect is ensured.
  • the wall which is opposite the step element and which delimits the cavity substantially perpendicular to the axis of the turbine does not itself have a step element.
  • actual labyrinth seals are problematic because they can limit the play function of the gap and adversely affect the mountability.
  • the step element or the plurality of segments, each of which assigns a step element to one guide vane is preferably located on the wall which is located downstream in the direction of flow of the hot gas in the secondary combustion chamber, i. usually on the platform, arranged.
  • the outer platform is fastened to the vane support via an intermediate ring, wherein a further wall region of the cavity adjoining radially to the wall region of the outer platform is formed by this intermediate ring, and furthermore preferably at the transition between the wall region of the platform and further wall region of the intermediate ring is formed a further step.
  • the cavity also extends between the guide blade carrier and the carrier element, that is, it is a cavity projecting deep into the structure.
  • said width B is reduced by the step (designed as a single step) by at least 20%, preferably by at least 30%. Under specific conditions, even a reduction of at least 40% may be desirable. Typically, reductions of up to 70% are desirable. Any further reduction is usually impractical and could also affect desired purge currents.
  • step element is arranged on the wall region of the outer platform and none on the opposite, preferably designed as a radially extending plane wall of the Wandungssegments.
  • Such an outer platform does not necessarily extend far radially outward.
  • this wall area, on which the step element is arranged then also not formed by the platform, but by the outside arranged intermediate ring or the vane support.
  • the width of the cavity expands radially outside the step element, preferably via a substantially perpendicular to the flow direction of the hot gas in the cavity extending stage substantially back to the original width B in the inlet region, further preferably radially outwardly followed by a second stage, the in turn rejuvenated.
  • the width B in the inlet region in the axial direction is in the range of 1-25 mm.
  • Fig. 1 in a) an axial section of the transition region between the radial
  • Fig. 2 in a) in a detailed view of a cavity with step element, in b) a
  • Fig. 3 is a schematic view in the radial direction of a segment of Cavity with circumferential step element and in b) a corresponding view with juxtaposed segments of step elements.
  • FIG. 1 a shows an axial section through the radially outer wall region of a gas turbine with secondary combustion chamber 1 in the transition from the secondary combustion chamber 1 to the first guide blade row 2 of the low-pressure turbine.
  • the radially inner boundary of the flow channel of the hot gases 3 is not shown.
  • the flow channel within the secondary combustion chamber 1 is formed radially on the outside by an outer wall segment 4. This is typically made of metal or ceramic, the metal typically being provided with a thermal protection layer.
  • This outer wall segment 4 is attached to the housing via a carrier element 5 and is usually acted upon from the rear with corresponding cooling air streams, which optionally additionally emerge through cooling air openings in Wandungssegment 4 in the hot air flow in the sense of film cooling.
  • Guide vanes are typically one-piece structures, which not only the actual vane, but molded thereon also include an inner platform and an outer platform 6.
  • the guide vanes can also be combined as assemblies of several vanes.
  • the outer platforms 6 form a circumferential ring, which tapers in the direction of flow.
  • the outer platforms 6 resp. these units of vanes and inner resp. outer platform 6 are attached to a so-called intermediate ring 7, which in turn is attached to the housing on a so-called guide vane carrier 8 of the low-pressure turbine.
  • a gap is formed which forms a cavity 9 which extends deeply into the housing components.
  • This cavity 9 is shown in more detail in Figure lb. Because of the bow wave already described at the leading edge of a respective vane results in particular at these radial positions, a high hot gas pressure at the inlet region of this cavity 9. Accordingly, there is a schematically indicated by the arrow 10 hot gas flow in this inlet region which, as shown schematically by the arrow 11, penetrates deep into the cavity.
  • the cavity 9 is to some extent formed on the downstream side (relative to the main flow direction of the hot gases 3) by a wall region 12 of the outer platform 6, followed by a wall region 13 of the intermediate ring 7 and still further radially outwardly followed by a wall region 14 of the guide blade carrier 8
  • these wall portions 12-14 are substantially aligned in a plane.
  • the further upstream in the flow direction and oppositely disposed wall, which limits the cavity 9, is radially formed on the inside first by the wall portion 15 of the outer wall segment 4 of the secondary combustion chamber, radially outwardly followed by the wall portion 16 of the support member 5 for the Wandungssemgent 4.
  • these wall portions 15, 16 in the constructions of the prior art.
  • the hot air flow 11 not only leads to unnecessarily high temperatures being reached in the cavity, but also leads, in particular, to oxidation problems at the wall regions 12-16. On the other hand, this gap is necessary for assembly reasons.
  • this gap or cavity 9 has a width B, which is shown in the contour illustration of the cavity 9 in FIG. 1c.
  • This width is typically in the range of 1-25 mm, ie in this area the gap is wide and correspondingly accessible for said hot air flow.
  • the hot gas passes through the entrance slit 17 and the circumferential projection 18 and forms behind this a hot air vortex 20 in the inlet region. Radially outside this vortex then flows the hot gas substantially unhindered in the radial direction and flows in accordance with high temperatures, ie with high oxidative effect, deep into the gap of the cavity 9 a.
  • FIG 2a shows a section analogous to Figure lb, which is now additionally formed with a step element 22 according to the invention.
  • This step element is designed as a circumferential rib arranged on the wall region 12 or integrally formed therewith, which provides a radially inward step directly in the flow direction of the hot gas behind the encircling projection.
  • this step element 22 extends in the radial direction approximately over a third or even half of the radial extent of the wall portion 12.
  • the opposite wall 15, however, except for the circumferential projection 18 in the entrance slit 17 designed flat and not also with a step element or with a corresponding formed corresponding groove.
  • the step element 22 to some extent forms a barrier to the hot gas flow and turbulence reduces the velocity of the hot gas.
  • leakage flows and purging air streams can then cool and protect the corresponding wall regions much more efficiently.
  • step element 22 is additionally molded onto the outer platform 6, but also the wall region located behind it is slightly milled out, resp. except that radially outside the step element 22, the width widens a little more than before and then forms a pronounced step 29 at the transition 23 to the wall portion 13. This level 29 results in additional turbulence and an extended additional barrier function.
  • FIG. 2c shows schematically how the flow conditions in such a structure are represented.
  • a first vortex 20 substantially behind the circumferential projection 18, but this is substantially reinforced by the entry step of the step element 22.
  • this vortex is much more formed than in Figure 1 and also unfolds a higher barrier effect.
  • a first vortex 24 forms in the region of the step element 22.
  • a second vortex 25 forms to some extent at the radially outer end of the step element in the widening region; these vortices 24, 25 lead to an additional barrier effect.
  • a further slight swirling favors an additional step 29 at the transition 23, the turbulence and leads to a further additional barrier function.
  • the temperature can already be massively reduced already in the area of the step element 22 but also radially outside of it by these measures, that lower pressures prevail and correspondingly the areas arranged in the area of the step element 22 and radially outside can be easily protected with cooling air.
  • FIG. 3 a shows how the step element 22 'can be formed circumferentially, ie in the form of a substantially circumferential ring about the axis of the low-pressure turbine.
  • the actually serious problems occur, as explained above, mainly at the leading edge of the respective guide blade 26.
  • Low-pressure turbine 23 Step transition from 12 to 13 outer cavity 24 first vortex
  • Wall area of 6 27 entry area of 9 adjacent to 9 28 first step on 22

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

L'invention concerne une turbine à gaz munie d'une chambre de combustion secondaire (1) et d'une première rangée d'aubes directrices (2), disposée directement en aval de cette dernière, d'une turbine basse pression. La chambre de combustion secondaire (1) est délimitée radialement vers l'extérieur par au moins un segment de paroi (4) extérieur fixé à un élément support (5), disposé radialement à l'extérieur. Le trajet de circulation des gaz chauds (3) dans la zone de la rangée d'aubes directrices (2) est délimité radialement vers l'extérieur par une plate-forme extérieure (6) fixée au moins indirectement à au moins un support d'aube directrice (8). Entre le segment de paroi (4) et la plate-forme extérieure (6) est formée une cavité (9) en forme de fente orientée sensiblement radialement présentant, dans la zone d'entrée, une largeur (B) dans la direction axiale située comprise entre 1 et 25 mm. La turbine à gaz selon l'invention est caractérisée en particulier en ce que, dans la zone d'entrée, est agencé au moins un élément à gradin (22, 22', 22'') qui réduit ladite largeur (B) d'au moins 10 % dans au moins un gradin (28) s'étendant dans la cavité (9) de manière sensiblement perpendiculaire au sens de circulation (11) du gaz chaud.
EP11716539A 2010-05-05 2011-04-26 Zone de transition pour une chambre de combustion secondaire d'une turbine à gaz Withdrawn EP2567071A1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
CH00691/10A CH703105A1 (de) 2010-05-05 2010-05-05 Gasturbine mit einer sekundärbrennkammer.
PCT/EP2011/056582 WO2011138193A1 (fr) 2010-05-05 2011-04-26 Zone de transition pour une chambre de combustion secondaire d'une turbine à gaz

Publications (1)

Publication Number Publication Date
EP2567071A1 true EP2567071A1 (fr) 2013-03-13

Family

ID=42340790

Family Applications (1)

Application Number Title Priority Date Filing Date
EP11716539A Withdrawn EP2567071A1 (fr) 2010-05-05 2011-04-26 Zone de transition pour une chambre de combustion secondaire d'une turbine à gaz

Country Status (9)

Country Link
US (1) US9097119B2 (fr)
EP (1) EP2567071A1 (fr)
JP (1) JP5826253B2 (fr)
KR (1) KR101645954B1 (fr)
CN (1) CN102884282B (fr)
CH (1) CH703105A1 (fr)
MX (1) MX2012012797A (fr)
RU (1) RU2540350C2 (fr)
WO (1) WO2011138193A1 (fr)

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US9243508B2 (en) * 2012-03-20 2016-01-26 General Electric Company System and method for recirculating a hot gas flowing through a gas turbine
JP5490191B2 (ja) 2012-07-19 2014-05-14 三菱重工業株式会社 ガスタービン
US10443736B2 (en) 2015-10-01 2019-10-15 United Technologies Corporation Expansion seal

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JPS62176448U (fr) 1986-04-30 1987-11-09
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Also Published As

Publication number Publication date
CN102884282A (zh) 2013-01-16
US20130055717A1 (en) 2013-03-07
US9097119B2 (en) 2015-08-04
RU2012152096A (ru) 2014-06-10
CH703105A1 (de) 2011-11-15
KR20130094184A (ko) 2013-08-23
MX2012012797A (es) 2013-01-17
WO2011138193A1 (fr) 2011-11-10
JP5826253B2 (ja) 2015-12-02
RU2540350C2 (ru) 2015-02-10
JP2013528738A (ja) 2013-07-11
KR101645954B1 (ko) 2016-08-12
CN102884282B (zh) 2015-07-29

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