EP1384950A2 - Annular combustion chamber for a gas turbine - Google Patents

Abstract

The ring-shaped combustion chamber (13), for a gas turbine (10), has burners (14,15) opening into it at the entry side aligned axially towards the outflow opening (33) at the other end. The chamber has cooled inner cladding segments (16,17) as a protection against the hot gases. The cladding segments are arranged axially in a number of successive sections. The cladding segments are cooled by convection, and the coolant is released from the downstream segment into the hot gas flow.

Description

TECHNICAL AREA

The present invention relates to the field of gas turbine technology. It relates to an annular combustion chamber for a gas turbine according to the Preamble of claim 1.

Such a combustion chamber as e.g. 3 is shown at Gas turbines in use for a long time.

STATE OF THE ART

In Fig. 3 is a sectional combustion chamber, a so-called EV combustion chamber (EV = Environmental), according to the state of the art played. The combustion chamber 26, which is part of a gas turbine, not shown and of which only the section above the turbine axis is reproduced extends in the longitudinal direction along the turbine axis in Flow direction (from right to left in Fig. 3). On the entry page (right Page in Fig. 3) is a number on a circular ring concentric to the turbine axis distributed by burners 27, which in the present case are known as so-called Double cone burners are designed according to EP 0321809. However, this is not imperative, and it goes without saying that those discussed here Combustion chambers can also be operated with other types of burners. The swirling fuel-air mixture emerging from the burners 27 burns with flame formation in the primary zone following the burner 27 30 and the resulting hot gases emerge from the combustion chamber 26 at one Combustion chamber outlet 31 out and into the subsequent turbine section, where it expand under work performance. Around the combustion chamber walls 29 in front of the hot ones Protecting gases are on the inside of the combustion chamber walls 29 special lining segments ("liner segments") 28 arranged and attached. The lining segments 28 are continuous in the axial direction and therefore as long as the interior of the combustion chamber 26. This has the advantage that the number of parts and the length of the leaky column is minimal.

However, the known configuration of the lining segments has the disadvantage that that the segments are comparatively long. This creates in terms of manufacturability and mechanical integrity problems. These problems will be even larger and possibly not solvable if for very large gas turbines correspondingly large combustion chambers with very long lining segments are needed.

PRESENTATION OF THE INVENTION

It is therefore an object of the invention to provide a combustion chamber which Avoids disadvantages of known combustion chambers described above and itself by simplifying production and assembly, as well as improved mechanical stability and mechanical and thermal resilience.

The object is achieved by the entirety of the features of claim 1. The essence of the invention is that in a combustion chamber the mentioned type, the lining segments in the axial direction in several successive arranged parts are divided. By dividing the individual Partial elements smaller, which simplifies their manufacture and the mechanical stability increased. At the same time, the assembly of the segments is simplified.

It has been found to be particularly favorable if the lining segments according to a preferred embodiment of the invention in two parts are divided when the liner segments are divided where the flow rate the hot gases is low, or if the liner segments are divided such that the lengths of the individual segment parts in axial direction are approximately the same.

The assembly can be further simplified if according to another embodiment the invention the lining segments attached to segment carriers are, and the segment carrier also divided into several parts in the axial direction are.

The lining segments are preferably convection-cooled.

The divided lining segments can be separately convection-cooled be through the downstream portions of the liner segments flowing cooling medium into the hot gas flow of the combustion chamber becomes.

But it is also conceivable that between the subdivided lining segments Connection channels are provided through which the convectively cooling cooling medium from one part of the lining segments to the other part of the lining segments can flow.

Further embodiments result from the dependent claims.

BRIEF EXPLANATION OF THE FIGURES

Fig. 1

einen Schnitt durch eine in einer Gasturbine angeordnete Brennkammer mit in axialer Richtung unterteilten Auskleidungssegmenten gemäss einem bevorzugten Ausführungsbeispiel der Erfindung;

Fig. 2

einen vergrösserten Ausschnitt aus der Darstellung der Fig. 1; und

Fig. 3

einen Schnitt durch eine ringförmige Brennkammer nach dem Stand der Technik.

The invention will be explained in more detail below on the basis of exemplary embodiments in connection with the drawing. Show it

Fig. 1

a section through a combustion chamber arranged in a gas turbine with lining segments subdivided in the axial direction according to a preferred embodiment of the invention;

Fig. 2

an enlarged section of the illustration of Fig. 1; and

Fig. 3

a section through an annular combustion chamber according to the prior art.

WAYS OF CARRYING OUT THE INVENTION

In Fig. 1 is a section through a combustion chamber arranged in a gas turbine with lining segments subdivided in the axial direction according to a preferred Embodiment of the invention reproduced. The gas turbine 10, from which is only shown a part lying above the turbine axis has one outer turbine housing 11 on which a filled with compressed air Surrounds plenum 12, in which the actual annular combustion chamber 13 is arranged is. The flow course takes place from right to left in FIG. 1. By in A head space of the combustion chamber 13 arranged burners 14, 15, which in two The fuel-air mixture enters the primary zone one above the other 32 blown into the combustion chamber 13 and burns there with the formation of flames. the resulting hot gases pass through the combustion chamber outlet 33 out of the combustion chamber 13 and into the subsequent turbine. The combustion chamber 13 is separated from the surrounding plenum by several segment carriers 18, .., 21 12 separated. On the inner walls of the segment carrier 18, .., 21 are in axial First and second lining segments 16 and 17 fastened in the direction one behind the other, with inner (in Fig. 1 lower) and outer (in Fig. 1 upper) lining segments are provided. The split liner segments 16, 17 have about the same (axial) length and are separated where the corresponding ones Butt segment carriers 19, 20 and 18, 21 together. The place where the butted divided liner segments 16, 17 (space 24 in Fig. 2), is where the flow rate of the hot gases is low. The split liner segments 16, 17 are convection cooled in the same way, as is already the case with the undivided lining segments is.

The division of the segment carriers 18, .., 21 ensures that the assembly is simplified. This applies in particular to the inner (lower) lining. If the inner lining is assembled from two parts, the Separation gap screwed along the entire length. The dividing line the segment carrier 18, 21 for the second lining segments 17 are for Bolt accessible so that a wedge is no longer needed.

The division of the lining segments according to the invention makes it possible to to realize larger combustion chambers without having correspondingly large ones Segments must be constructed. This way you can look at already proven ones Use segment sizes. The invention also makes it possible in different Gas turbines have the same burners 14, 15 and first liner segments 16 to use. Adapted to different turbine inlet geometries then only the combustion chamber outlet 33 with the second lining segments 17 and their segment carriers 18, 21.

The configuration of the liner segments 16, 17 is the same as for the EV and SEV combustion chambers of the known gas turbines of the applicant of the type GT24B and GT26B (see the article by D. K. Mukherjee "State-of-the-art gas turbines - a brief update ", ABB review 2/1997, pp. 4-14 (1997)). A specialty is the provision of connecting channels 22, 23 (Fig. 1 and Fig. 2) between the second liner segments 17 and the first liner segments 16. Through these connecting channels 22, 23 for the convective cooling of the Lining segments 16, 17 used cooling air from the second lining segments 17 flow into the first lining segments 16 and there to Contribute cooling. The cooling system of the second liner segments 17 is operated only with a part of the total cooling mass flow to the flow velocities to avoid pressure drops in the connection channels 22, 23 to keep small. For cooling the first lining segments 16 an additional partial flow 25 is required (FIG. 2). The transition area between the inner second and first liner segments 17 and 16 is in Fig. 2 shown enlarged.

But it is also conceivable to dispense with the connecting channels 22, 23 and Cooling systems of the first and second liner segments 16, 17 separately train. The cooling air from the second liner segments 17 is then released into the hot gas flow. The second liner segments 17 are significantly shorter and are optimized for minimal cooling air consumption. The advantage of separate cooling is that on the manufacturing side elaborate connecting channels 22, 23 can be dispensed with, and that air for Influencing the hot gas temperature distribution and for cooling the gap between the combustion chamber and turbine is available. This is bought Benefit from a reduced air mass flow in the burner and a low one Height of the cooling channels in the second lining segments 17.

LIST OF REFERENCE NUMBERS

10

gas turbine

11

outer turbine housing

12

plenum

13.26

Combustion chamber (ring-shaped)

14,15,27

burner

16.17

liner segment

18, .., 21

segment carrier

22.23

connecting channel

24

gap

25

partial flow

28

liner segment

29

combustion chamber wall

30.32

primary zone

31.33

combustor exit

Claims (10)

Annular combustion chamber (13) for a gas turbine (10), into which combustion chamber (13) open on an inlet side burners (14, 15), and which combustion chamber (13) extends in the axial direction from the inlet side to an outlet side (33) and on the inside for protection against the hot gases is lined with cooled lining segments (16, 17), characterized in that the lining segments (16, 17) are divided in the axial direction into several parts (16, 17) arranged one behind the other. Combustion chamber according to claim 1, characterized in that the lining segments (16, 17) are divided into two parts (16, 17). Combustion chamber according to claim 2, characterized in that the lining segments (16, 17) are divided where the flow rate of the hot gases is low. Combustion chamber according to claim 3, characterized in that the lining segments (16, 17) are subdivided in such a way that the lengths of the individual segment parts (16, 17) are approximately the same in the axial direction. Combustion chamber according to one of claims 1 to 4, characterized in that the lining segments (16, 17) are attached to segment carriers (18, .., 21), and that the segment carriers (18, .., 21) are also in the axial direction in several parts (18, .., 21) are divided. Combustion chamber according to one of claims 1 to 5, characterized in that the lining segments (16, 17) are convection-cooled. Combustion chamber according to claim 6, characterized in that the divided lining segments (16, 17) are separately convection-cooled. Combustion chamber according to claim 7, characterized in that the cooling medium flowing through the downstream parts (17) of the lining segments is discharged into the hot gas flow of the combustion chamber (13). Combustion chamber according to claim 6, characterized in that between the divided lining segments (16, 17) there are connection channels (22, 23) through which the convectively cooling cooling medium from one part (17) of the lining segments into the other part (16) of the lining segments can flow. Combustion chamber according to one of claims 6 to 8, characterized in that the downstream parts (17) of the lining segments are cooled only by a part of the mass flow provided overall for cooling the lining segments.

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