EP1384950B1 - Annular combustion chamber for a gas turbine - Google Patents

Description

TECHNICAL AREA

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

Such a combustion chamber, such as in the Fig. 3 is in gas turbines for a long time in use.

STATE OF THE ART

In Fig. 3 is a sectional view of an annular combustion chamber, a so-called EV (EV) combustion chamber, reproduced according to the prior art. The combustion chamber 26, the part of a gas turbine, not shown is reproduced and of which only the section lying above the turbine axis, extends in the longitudinal direction along the turbine axis in the flow direction (in Fig. 3 from right to left). On the entrance side (right side in Fig. 3 ) is arranged distributed on a concentric to the turbine axis annulus a number of burners 27, which in the present case as a so-called double-cone burner according to EP 0321809 are formed. However, this is not a mandatory requirement, and it goes without saying that the combustion chambers discussed here can also be operated with other burner variants. The emerging from the burners 27, fluidized air-fuel mixture burns with flame in the following on the burner 27 primary zone 30 and the resulting hot gases emerge from the combustion chamber 26 at a Brennkammeraustritt 31 and in the subsequent turbine part, where it Expand work performance. In order to protect the combustion chamber walls 29 from the hot gases, special lining segments 28 are arranged and fastened on the inside of the combustion chamber walls 29. 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 gap is minimal.

A disadvantage of the known configuration of the lining segments, however, is that the segments are comparatively long. This creates problems in terms of manufacturability and mechanical integrity. These problems are even greater and may not be solved if correspondingly large combustion chambers with very long lining segments are required for very large gas turbines.

In addition, however, it is known per se in the field of gas turbine technology to equip combustors with a plurality of smaller liner segments. However, the smaller these segments are dimensioned, the more obvious is the problem of reliably sealing the existing column, which in turn must be compensated by other expensive measures. Insufficient thermal insulation, for example due to penetrating Hot gases, affects the function of the gas turbine significantly and not least has a negative impact on their service life.

US 5363643 deals with the problem of the use of lining segments based on ceramic materials in the combustion chamber of an aircraft engine. The high heat resistance of these materials suggests their use. However, their thermal expansion behavior, which differs significantly from that of the metallic support structure, results in the aforementioned considerable risks during operation of the engine. By arranging a plurality of rectangular lining segments with special structural features to overcome these disadvantages.

In the solution according to EP 1363075 The numerous, essentially at least approximately rectangular, lining segments are equipped with film cooling holes.

US 4446693 also describes features of the wall structure of the combustion chamber of an aircraft engine, which, however, is characterized by arranging a plurality of overlapping lining segments with outlet gaps for cooling air in the overlapping region.

PRESENTATION OF THE INVENTION

It is therefore an object of the invention to provide a combustion chamber which avoids the disadvantages of known combustion chambers described above and is characterized by a simplification of the production and assembly and improved mechanical stability.

The object is solved by the entirety of the features of claim 1. The essence of the invention is that in a combustion chamber of the type mentioned, the lining segments are divided in the axial direction into a plurality of successively arranged parts, and are secured to the segment carriers, which are also divided in the axial direction into several parts.

By dividing the lining segments, the individual sub-elements are smaller, which simplifies their production and increases the mechanical stability, and by the segment carrier to which the lining segments are attached, are also divided in the axial direction into several parts, simplifies the installation of the same time segments.

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

The lining segments are preferably convection-cooled.

In this case, the divided liner segments may be separately convection cooled, wherein the flowing through the downstream parts of the lining segments cooling medium is discharged into the hot gas flow of the combustion chamber.

However, it is also conceivable that connection channels are provided between the subdivided lining segments, through which the convectively cooling cooling medium flows from one part of the lining segments into the other part of the lining segments.

Further embodiments emerge 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 with reference to embodiments in conjunction with the drawings. Show it

Fig. 1

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

Fig. 2

an enlarged excerpt from the depiction of the Fig. 1 ; and

Fig. 3

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

WAYS FOR CARRYING OUT THE INVENTION

In Fig. 1 is a section through a arranged in a gas turbine combustion chamber with divided in the axial direction lining segments according to a preferred embodiment of the invention reproduced. The gas turbine 10, of which only one lying above the turbine axis part is shown, has an outer turbine housing 11 which surrounds a filled with compressed air plenum 12, in which the actual annular combustion chamber 13 is arranged. The flow takes place in Fig. 1 from right to left. By arranged in a headspace of the combustion chamber 13 burners 14, 15, which lie in two rows one above the other, the fuel-air mixture is blown into the primary zone 32 of the combustion chamber 13 and burns there to form flames. The resulting hot gases exit through the combustion chamber outlet 33 from the combustion chamber 13 and into the subsequent turbine. The combustion chamber 13 is separated by a plurality of segment carrier 18, .., 21 from the surrounding plenum 12. On the inner walls of the segment carrier 18, .., 21 in the axial direction one behind the other first and second lining segments 16 and 17 are fixed, wherein each inner (in Fig. 1 lower) and outer (in Fig. 1 upper) lining segments are provided. The split liner segments 16, 17 have approximately the same (axial) length and are separated there, where the associated Segment carrier 19, 20 and 18, 21 abut. The location where the split liner segments 16, 17 abut (gap 24 in FIG Fig. 2 ), where the flow velocity of the hot gases is low. The split liner segments 16, 17 are convection cooled in the same manner as is already the case with the undivided liner segments.

By the division of the segment carrier 18, .., 21 is achieved that the assembly is simplified. This is especially true for the inner (lower) liner. When the inner lining is assembled from two parts, the separation gap can be screwed down the entire length. The dividing line of the segment carriers 18, 21 for the second lining segments 17 is accessible to threaded bolts, so that a wedge is no longer needed.

The inventive division of the lining segments and the segment carrier, it is possible to realize larger combustion chambers without correspondingly large segments must be constructed. In this way one can fall back on already proven segment sizes. The invention also makes it possible to use the same burners 14, 15 and first liner segments 16 in different gas turbines. Adapted to different turbine inlet geometries then only the combustion chamber outlet 33 with the second lining segments 17 and their segment carriers 18, 21st

The configuration of the liner segments 16, 17 is the same as in the EV and SEV combustors of the prior art GT24B and GT26B type gas turbine engines (see the article of US Pat DK Mukherjee "State-of-the-art gas turbines - a letter update", ABB Review 2/1997, pp. 4-14 (1997 )). A special feature is the provision of connection channels 22, 23 (FIG. Fig. 1 and Fig. 2 ) between the second lining segments 17 and the first lining segments 16. Through these connecting channels 22, 23, the cooling air used for the convective cooling of the lining segments 16, 17 flow from the second lining segments 17 into the first lining segments 16 and there to Contribute cooling. The cooling system of the second liner segments 17 is operated with only a portion of the total cooling mass flow to keep the flow rates to avoid pressure drops in the connection channels 22, 23 small. For the cooling of 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.

However, it is also conceivable to dispense with the connecting channels 22, 23 and form the cooling systems of the first and second lining segments 16, 17 separately. The cooling air from the second liner segments 17 is then discharged into the hot gas stream. The second lining segments 17 are significantly shorter and are optimized for a minimum cooling air consumption. The advantage of the separate cooling is that it is possible to dispense with the production-technically complicated connection channels 22, 23, and that air is available for influencing the hot gas temperature distribution and for cooling the gap between the combustion chamber and the turbine. This advantage is paid for by a reduced air mass flow in the burner and a small 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 (9)

1. An annular combustor (13) for a gas turbine (10), into which combustor (13) burners (14, 15) open on an inlet side, and which combustor (13) extends in the axial direction from the inlet side to an outlet side (33) and is lined on the insides with cooled liner segments (16, 17) for protection from the hot gases, and these liner segments (16, 17) are subdivided in the axial direction into several parts (16, 17) arranged one behind the other, characterized in that the liner segments (16, 17) are fastened to segment carriers (18,...,21), and these segment carriers (18,...,21) are likewise subdivided in the axial direction into several parts (18, ...,21) .
2. The combustor as claimed in claim 1, characterized in that the liner segments (16, 17) are subdivided into two parts (16, 17).
3. The combustor as claimed in claim 2, characterized in that the liner segments (16, 17) are subdivided where the flow velocity of the hot gases is low.
4. The combustor as claimed in claim 3, characterized in that the liner segments (16, 17) are subdivided in such a way that the lengths of the individual segment parts (16, 17) in the axial direction are approximately the same.
5. The combustor as claimed in one of claims 1 to 4, characterized in that the liner segments (16, 17) are convection-cooled.
6. The combustor as claimed in claim 5, characterized in that the subdivided liner segments (16, 17) are convection-cooled separately.
7. The combustor as claimed in claim 6, characterized in that the cooling medium flowing through those parts (17) of the liner segments which are situated downstream is released into the hot-gas flow of the combustor (13).
8. The combustor as claimed in claim 5, characterized in that transition channels (22, 23) are provided between the subdivided liner segments (16, 17), through which transition channels (22, 23) the convectively cooling cooling medium flows from one part (17) of the liner segments into the other part (16) of the liner segments.
9. The combustor as claimed in one of claims 5 to 7, characterized in that those parts (17) of the liner segments which are located downstream are cooled only by part of the mass flow provided overall for the cooling of the liner segments.

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