EP2567071A1

EP2567071A1 - Transition region for a secondary combustion chamber of a gas turbine

Info

Publication number: EP2567071A1
Application number: EP11716539A
Authority: EP
Inventors: Thomas Heinz-Schwarzmaier; Marc Widmer; Selma Zahirovic; Paul Marlow
Original assignee: Alstom Technology AG
Current assignee: Ansaldo Energia IP UK Ltd
Priority date: 2010-05-05
Filing date: 2011-04-26
Publication date: 2013-03-13
Also published as: MX2012012797A; RU2012152096A; RU2540350C2; JP2013528738A; KR101645954B1; US9097119B2; CN102884282B; US20130055717A1; WO2011138193A1; KR20130094184A; CN102884282A; CH703105A1; JP5826253B2

Abstract

The invention relates to a gas turbine having a secondary combustion chamber (1) and a first guide vane row (2) of a low pressure turbine, arranged directly downstream thereof, wherein the radial exterior delimitation of the secondary combustion chamber (1) is formed by at least one outer wall segment (4), which is fastened to at least one carrier element (5), arranged radially on the outside, wherein the flow path of the hot gases (3) is delimited radially on the outside by an outer platform (6) in the region of the guide vane row (2) which is fastened at least indirectly to at least one guide vane carrier (8), and wherein there is a gap-shaped cavity (9), running substantially radially and having a width (B) in the inlet region in the axial direction in the range of 1-25 mm, between the wall segment (4) and the outer platform (6). The gas turbine according to the invention is in particular characterized in that at least one step element (22, 22', 22'') is arranged in the inlet region, which reduces the cited width (B) by at least 10% in at least one step (28), running substantially perpendicularly to the flow direction (11) of the hot gas in the cavity (9).

Description

TITLE

TRANSITION AREA FOR A SECONDARY COMBUSTION CHAMBER ONE

GAS TURBINE

TECHNICAL AREA

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.

STATE OF THE ART

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. Typically, 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. On the opposite, closer to the axis inside there is a correspondingly formed inner boundary, an inner wall of corresponding materials. 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.

Accordingly, 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. For assembly reasons and due to the different mechanical and thermal loads on the components of the secondary combustion chamber and turbine, 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. At the leading edge of the vanes, a so-called bow wave or a so-called "horse shoe vortex" is formed, which causes hot combustion air in the wall area to be pressed into this cavity and penetrates deep into it. This can give rise to problems in connection with overheating but also with oxidation of the corresponding surfaces. From US 2009/0293488 it is known that it is possible to close this transition region substantially by a very small gap and additionally provide specific structures which ensure optimum cooling of the wall regions in this area. The problem with this approach, however, is that the required clearance between the combustion chamber module and the turbine is not readily ensured by the correspondingly small gap size. PRESENTATION OF THE INVENTION

Here, the present invention intervenes and goes a completely different way than the prior art. In particular, no attempt should be made to close the gap, as this results in the above-mentioned problems. On the contrary, 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.

Specifically, according to the present invention relates to 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.

According to the invention, 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%.

Through 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.

According to a first preferred embodiment, 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.

Particularly problematic are the areas which are arranged radially directly outside the respective leading edge of each vane. In particular in these areas, the turbulence caused by the bow wave penetrates particularly strongly into the cavity. The intervening areas, on the other hand, are less affected. Accordingly, it is also possible according to a further preferred embodiment, 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). Preferably, 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.

The 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. Preferably, 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. Furthermore, 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.

In general, it is preferred that 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. In other words, in the present invention, it is preferably not a matter of providing a labyrinth seal in the classical sense, in which the flow path is somewhat meander-shaped, but rather providing a step element on only one of the two opposite walls of the cavity. In fact, actual labyrinth seals are problematic because they can limit the play function of the gap and adversely affect the mountability.

In general, 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.

According to a further preferred embodiment, 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.

Preferably, the cavity also extends between the guide blade carrier and the carrier element, that is, it is a cavity projecting deep into the structure.

According to a further preferred embodiment of the invention, 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.

As already mentioned above, it is preferred if only one 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. In this case, 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.

Preferably, 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.

Preferably, the width B in the inlet region in the axial direction is in the range of 1-25 mm.

It is possible that immediately adjacent to the entrance gap to the cavity on the wall of the outer wall segment, a circumferential, the inlet gap locally tapered projection is formed.

Further embodiments are given in the dependent claims.

BRIEF DESCRIPTION OF THE DRAWINGS

Preferred embodiments of the invention will be described below with reference to the drawings, which are given by way of illustration only and are not to be interpreted as limiting. In the drawings show:

Fig. 1 in a) an axial section of the transition region between the radial

Outside wall of the secondary combustion chamber and the outer platform of the first row of vanes of the low-pressure turbine, wherein the corresponding vane is not shown, in b) a detailed view of the section according to a) with marked hot air flows in the cavity, in c) a contour representation of the cavity and in d) a schematic representation of the flow conditions in the inlet region of the cavity;

Fig. 2 in a) in a detailed view of a cavity with step element, in b) a

Contour representation of a cavity with step element and behind set back wall of the outer platform, and in c) a schematic

Representation of the flow conditions in the inlet region of the cavity with

Step element; and

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.

DESCRIPTION OF PREFERRED EMBODIMENTS

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.

In the direction of flow of the hot gas 3 downstream of the secondary combustion chamber follows the first row of vanes 2. 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 platforms which cover a segment viewed around the turbine axis, when a series of such vane elements are arranged around the circumference of a gas turbine, form not only the attachment of the vanes, but simultaneously also the radially outer boundary in the case of the outer platform 6 the flow path for the hot gas and in the case of the inner platform, the inner boundary of this flow path. In other words, 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.

Between the wall elements 4 of the secondary combustion chamber 1 and the outer platform 6 of the first row of guide vanes 2 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. In this case, 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 In the prior art designs, 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. Here too 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.

In the inlet region 27, 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. Immediately at the entrance slit 17 in this cavity 9, there is a from the outer wall segment 4 at the radially leading edge in the flow direction of the hot gas extending circumferential projection 18, which respectively the entry resp. the front entrance gap slightly reduced. Behind it, however, the entrance slit widens again to the stated width B. In the case of such a slit without special measures, a flow pattern then results, as shown schematically in FIG. 1 d. 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.

Figure 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. Typically, 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. Accordingly, the step element 22 to some extent forms a barrier to the hot gas flow and turbulence reduces the velocity of the hot gas. Correspondingly, leakage flows and purging air streams can then cool and protect the corresponding wall regions much more efficiently. Both the inlet gap facing the step of the step element 22 and the radially outer step behind the step element 22, where the cavity widens again, leads to a vortex formation.

In the contour illustration according to FIG. 2b, not only the 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. As before, there is a first vortex 20 substantially behind the circumferential projection 18, but this is substantially reinforced by the entry step of the step element 22. In other words, this vortex is much more formed than in Figure 1 and also unfolds a higher barrier effect. In addition, 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. ever After detail geometry and purging air flow, a further slight swirling favors an additional step 29 at the transition 23, the turbulence and leads to a further additional barrier function. Looking now at the temperatures, it can be seen that 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. Correspondingly, it may also be sufficient, as shown in FIG. 3b, if, to a certain extent, radially outside each guide vane and coordinated with its leading edge, only one segment 22 "of such a step element is arranged in order to produce the effect according to the invention.

LIST OF REFERENCE NUMBERS

Secondary combustion chamber adjacent to 9

Guide vane row 17 entry slit in 9

Hot gas ström 18 circumferential projection outer wall segment

of 1 20 vertebrae in the entry area

Carrier element for 4

outer platform of 26 22 step element

Intermediate ring 22 'step element, circumferential

Guide vane carrier of 22 "step element, segmental

Low-pressure turbine 23 Step transition from 12 to 13 outer cavity 24 first vortex

Hot gas stream entry into 9 25 second vortex

Hot gas flows into 9 26 vane

Wall area of 6 27 entry area of 9 adjacent to 9 28 first step on 22

Wall area of 7 29 level at 23

adjacent to 9

Wall area of 8 P pitch

adjacent to 9 B width in the entry area

Wall area of 4

adjacent to 9

Wall area of 5

Claims

A gas turbine with a secondary combustion chamber (1) and a first Leitschaufelreihe 2 of a low-pressure turbine disposed immediately downstream thereof, wherein the radially outer boundary of the secondary combustion chamber (1) by at least one outer Wandungssegment (4) is formed, which at least one radially outwardly disposed support member (5), wherein the flow path of the hot gases (3) in the region of the guide vane row is bounded radially on the outside by an outer platform (6) which is at least indirectly fixed to at least one vane carrier (8), and between the wall segment (4 ) and the outer platform (6) a substantially radially extending gap-shaped cavity (9) having a width (B) in the inlet region in the axial direction in the range of 1-25 mm is present, characterized in that

at least one step element (22, 22 ', 22 ") is arranged in the inlet region, which surrounds said width (B) in at least one step (28) extending substantially perpendicular to the flow direction (11) of the hot gas in the cavity (9) at least 10% reduced.

2. Gas turbine according to claim 1, characterized in that the step element (22, 22 ') is formed circumferentially with respect to the axis of the turbine.

3. Gas turbine according to one of the preceding claims, characterized in that the step element (22, 22 ") is formed circumferentially segmentally and each vane (26) radially outside such a segment (22") is assigned, preferably all segments have a length in Circulation direction relative to the orbiting distance (p) of the guide vanes (26) of 30-50% of the orbiting distance (p) have.

4. Gas turbine according to one of the preceding claims, characterized in that the step element (22) in the form of a to the cavity (9) adjacent wall region (12) of the outer platform (6) patched or molded is formed substantially in axial cross-section rectangular rib, which preferably has a length in the radial direction in the range of 10-100 mm, in particular preferably in the range of 20-50 mm, preferably in combination with a radially arranged outside in this wall region (12) formed recess of an equal or greater length, the radially outer side end is formed by a further step (29).

5. Gas turbine according to one of the preceding claims, characterized in that the outer platform (6) via an intermediate ring (7) is fixed to the guide blade carrier (8), wherein another, radially to the wall region (12) of the outer platform (6) adjacent wall region (13) of the cavity (9) is formed by this intermediate ring (7), wherein preferably at the transition (23) between the wall region (12) of the platform (6) and the further wall region (13) of the intermediate ring (7) further stage (29) is formed.

6. Gas turbine according to one of the preceding claims, characterized in that the cavity (9) also extends between the guide blade carrier (8) and the carrier element (5).

7. Gas turbine according to one of the preceding claims, characterized in that said width (B) through the step (28) by at least 20%, preferably reduced by at least 30%.

8. Gas turbine according to one of the preceding claims, characterized in that both on the outer platform (6) and on the wall (15) of the Wandungssegment (4) and / or the wall region (5) at least one step element (22) is arranged.

9. Gas turbine according to one of the preceding claims, characterized in that only one step element (22) on the wall region (12) of the outer platform (6) is arranged and none at the opposite, preferably designed as a radially extending plane wall (15) of the wall segment (4).

10. Gas turbine according to one of the preceding claims, characterized in that the width substantially radially outside of the step element (22), preferably via a substantially perpendicular to the flow direction (11) of the hot gas in the cavity (9) extending stage substantially back to the original width (B) extended in the entry area.

11. Gas turbine according to one of the preceding claims, characterized in that the width (B) in the inlet region in the axial direction in the range of 2-15 mm.

12. Gas turbine according to one of the preceding claims, characterized in that immediately at the inlet gap (17) to the cavity (9) on the wall (15) of the outer wall segment (4) has a circumferential, the inlet gap locally tapered projection (18). is trained.