EP0597137B1 - Combustion chamber for gas turbine - Google Patents

Description

Technical field

The invention relates to a gas turbine combustor with an annular combustion chamber according to the preamble of claim 1.

State of the art

Gas turbine combustion chambers with air-cooled flame tubes are known, for example from US 4,077,205 or US 3,978,662. The flame tube is essentially constructed from wall parts which overlap in the turbine axial direction. On their side facing away from the combustion chamber, the wall parts each have a plurality of inlet openings distributed over the circumference, via which air is introduced into a distribution chamber arranged in the flame tube and communicating with the combustion chamber. In the cooling system there, the respective flame tube has a lip which extends over the slot through which the cooling air film emerges. This cooling air film should adhere to the wall of the flame tube in order to form a cooling barrier layer for it.

The so-called "lean premix combustion" has recently become established for the low-pollutant combustion of a gaseous or liquid fuel. The fuel and the combustion air are premixed as evenly as possible and only then fed to the flame. If this is done with a high excess of air, as is customary in gas turbine plants, relatively low flame temperatures arise, which in turn leads to the desired, low formation of nitrogen oxides.

The known gas turbine combustion chambers mentioned above now have the disadvantage that the air consumption for cooling purposes is much too high and that as a result of the infeed of the Cooling air into the flame tube interior downstream of the flame, this air is not available to the actual combustion process. As a result, the combustion chamber cannot be operated with the required high excess air ratio.

A gas turbine combustion chamber of the type mentioned at the beginning with an annular combustion chamber is known from GB-A-2074308. The segments forming the wall of the combustion chamber are suspended in a lattice-shaped frame. They consist of an inner and an outer wall. They are supplied with cooling air through openings on the outer wall.

A pot combustion chamber is known from US Pat. No. 4,288,980, in which the walls of the combustion chamber do not extend to the inlet of the gas turbine. In such pot combustion chambers there is an extremely complex hot gas housing between their outlet and turbine inlet, in which a transition from the circular combustion chamber cross section to the annular gas turbine cross section has to be made. The entire compressed combustion air is led through an annular space between walls around the hot gas housing and introduced into the combustion chamber. There is no separate and independent cooling of the primary and secondary zones. Instead, the entire combustion air flows into the cooling tubes of the primary zone after flowing through the ring channel and after cooling the wall there. The same air thus cools both the walls of the secondary zone and then the wall of the primary zone formed by the tubes. The outlet end of the flame tube is thus not in direct communication with the combustion chamber inlet, but opens into the tube via a flange for convective cooling there. This has the disadvantage that, on the one hand, the flow through the tubing is associated with a loss of pressure, and, on the other hand, that the highly stressed primary zone is subjected to cooling air which is already preheated.

Presentation of the invention

The invention has for its object to minimize the cooling air consumption in a gas turbine combustion chamber of the type mentioned in order to reduce the emission of NO _x .

According to the invention, this object is achieved with the features of the claims.

The advantage of the invention can be seen, inter alia, in the fact that with the new measure, as a result of the separate crossflow of the two cooling air flows, their pressure losses can be kept low. Finally, the entire cooling air is fed into the combustion process after cooling.

It is particularly expedient if, on the one hand, radial openings communicating with the collecting space for the supply of the segment cooling air and, on the other hand, axial channels communicating with the burner inlet are arranged for the joint removal of the segment cooling air and the cooling air acting on the secondary zone. In addition to its supporting function, the segment carrier also takes over the channeling of all cooling air flows. Since the carrier is usually a cast piece, the required openings can be made in the simplest way, which makes additional air lines unnecessary.

If premixing burners of the double-cone type are used as burners, as are known for example from EP-B1-321809, two such burners are generally arranged radially one above the other on a front segment. On the front segments assembled to form a circular ring, the burners of adjacent front segments are each radially offset for reasons of space. This means that every second burner is arranged closer to the segments in the circumferential direction than the immediately adjacent ones. If now in The number of cooling segments lined up corresponds to the number of front segments and if the number of air supply openings and discharge channels in the segment carrier also corresponds to the number of cooling segments in the circumferential direction, then you have a simple tool in hand, for example by differently dimensioning the inflow or outflow bores Dosing air supply to the cooling segments according to their thermal load.

Brief description of the drawing

In the drawing, an embodiment of the invention is shown using a single-shaft gas turbine with axial flow.

Fig. 1

einen Teillängsschnitt der Gasturbine;

Fig. 2

ein vergrösserter Ausschnitt der Primärzone der Brennkammer;

Fig. 3

einen Teilquerschnitt durch die Primärzone der Brennkammer nach Linie 3-3 in Fig. 2;

Show it:

Fig. 1

a partial longitudinal section of the gas turbine;

Fig. 2

an enlarged section of the primary zone of the combustion chamber;

Fig. 3

a partial cross section through the primary zone of the combustion chamber according to line 3-3 in Fig. 2;

Only the elements essential for understanding the invention are shown. The system does not show, for example, the complete exhaust pipe with chimney and the inlet parts of the compressor part. The direction of flow of the work equipment is indicated by arrows.

Way of carrying out the invention

The system, of which only the half lying above the machine axis 10 is shown in FIG. 1, essentially consists on the gas turbine side (1) of the rotor 11 bladed with rotor blades and the blade carrier 12 equipped with guide blades. The blade carrier 12 is over projections hooked into corresponding receptacles in the turbine housing 13. The exhaust housing 14 is flanged to the turbine housing 13, which essentially consists of a hub-side, annular inner part 16 and an annular outer part 17, which delimit the diffuser 19. Both elements 16 and 17 are usually half-shells with an axial parting plane. They are connected to one another by a plurality of radial flow ribs 18, which are arranged evenly distributed over the circumference. The outlet-side mounting of the turbomachine is arranged in the cavity within the inner part 16, the rotor 11 being located in a support bearing 21.

The turbine housing 13 and the blade carrier 12 are provided with a horizontal parting plane (not shown) lying in the machine axis 10. The upper and lower halves of the turbine housing and of the blade carrier, which are generally provided with flanges, are screwed together.

In the illustrated case, the turbine housing 13 also includes the collecting space 15 for the compressed combustion air. Part of the combustion air passes from this collecting space into the annular combustion chamber 3, which in turn enters the turbine inlet, i.e. flows upstream of the first guide row. The compressed air arrives in the collecting space from the diffuser 22 of the compressor 2. Only the last three stages of the latter are shown. The blading of the compressor and the turbine sit on the common shaft 11, the central axis of which represents the longitudinal axis 10 of the gas turbine unit.

The shaft part located between the turbine and the compressor is designed as a drum 23. The entire axial extent of this drum is surrounded by a drum cover 24, which is fastened to the diffuser outer housing of the compressor via ribs (not shown). This drum cover forms the cover band for the compressor Buckets of the last compressor guide row. On the turbine side, the drum cover, together with the end face of the turbine rotor, delimits a radially running wheel side space. This space forms the outlet-side end of an annular channel 25 which, starting from the hub behind the last row of compressor runs, runs between the drum cover and the drum. The entire rotor-side cooling air is introduced into this ring duct.

The combustion chamber 3 is equipped at its head end with premix burners 20, as are known for example from EP-B1-321 809. Such a premix burner, shown only schematically in FIG. 2, is a so-called double-cone burner. It essentially consists of two hollow, conical partial bodies 26, 27 which are nested one inside the other in the direction of flow. The respective central axes of the two partial bodies are offset from one another. The adjacent walls of the two partial bodies in their longitudinal extent form tangential slots 28 for the combustion air, which in this way reaches the interior of the burner. A fuel nozzle 29 for liquid fuel is arranged there. The fuel is injected into the hollow cone at an acute angle. The resulting conical liquid fuel profile is enclosed by the combustion air flowing in tangentially. The concentration of the fuel is continuously reduced in the axial direction due to the mixing with the combustion air. The burner can also be operated with gaseous fuel. For this purpose, gas inflow openings distributed in the longitudinal direction are provided in the region of the tangential slots in the walls of the two partial bodies. In gas operation, the mixture formation with the combustion air thus begins in the zone of the inlet slots 28. It goes without saying that mixed operation with both types of fuel is also possible in this way. At the burner outlet, a fuel concentration that is as homogeneous as possible is established over the applied annular cross section. A defined one is created at the burner outlet dome-shaped backflow zone, at the tip of which the ignition takes place.

Part of the compressed combustion air enters the burner from the collecting space 15 through a perforated cover 30 in the direction of the arrow. On the occasion of the combustion, the combustion gases reach very high temperatures, which places special demands on the combustion chamber walls to be cooled. This applies all the more when so-called low NO _x burners, for example the premix burners used here, are used, which require relatively modest amounts of cooling air. The annular combustion chamber extends downstream of the burner orifices up to the turbine inlet. It is limited both inside and outside by walls to be cooled, which are usually designed as self-supporting structures.

So far, ring combustion chambers for gas turbines are known.

The present combustion chamber is equipped with 72 of the said burners 20. 3, which shows a quarter-circle section, shows the arrangement thereof. Two burners are arranged radially one above the other on a front segment 31. 36 of these adjacent front segments form a closed circular ring, which in this way forms a heat shield. The two burners from adjacent front segments are each radially offset. This means that the radially outer burner of every second front segment directly adjoins the outer ring wall of the combustion chamber, as can also be seen in FIG. 2. The radially inner burners of the other front segments are therefore arranged in the immediate vicinity of the inner ring wall. This results in an uneven thermal load on the corresponding ring walls over the circumference.

The interior of the combustion chamber is now divided into two zones, the walls of which are cooled in different ways.

A secondary zone 32 lying downstream and opening into the turbine inlet is delimited by a double-walled flame tube. It consists both on its inner ring 33 and on its outer ring 34 from a flangeless, welded sheet metal construction, which is held together by spacers, not shown. Both rings 33 and 34 are open at their turbine end and form the entry for the cooling air there.

The annular space 35 between the double wall of the outer ring 34 draws the air directly from the collecting space 15, as can be seen in FIG. 1. With efficient convection cooling, the air flows in the counterflow to the combustion chamber flow in the direction of the primary zone 36.

The annular space 37 between the double wall of the inner ring 33 is supplied with air from a hub diffuser 38. This hub diffuser, which connects to the compressor diffuser 22, is delimited on the one hand by the drum cover 24 and on the other hand by an annular shell 39. The latter is connected to the drum cover 24 via ribs (not shown). In this annular space 37, too, the air flows in the counterflow to the combustion chamber flow in the direction of the primary zone 36.

The cooling of the highly stressed primary zone walls is now carried out according to the invention by means of individually cooled cooling segments 40. These cooling segments lined up in the circumferential direction and in the axial direction form their flow-limiting wall over the entire axial extent of the primary zone 36. The individual cooling has the advantage of a low pressure drop and the cooling can be adapted to local conditions.

The thermally highly stressed cooling segments 40 consist of a high-temperature, precision cast alloy. They are in the circumferential direction, each with two feet with supporting teeth 42 suspended in corresponding grooves in a supporting structure, similar to how guide vane feet are fastened in blade carriers. Similarly to blade carriers, this support structure, hereinafter referred to as segment carrier 43, consists of two cast half-shells with a horizontal parting plane and not shown claws with which it is supported in the turbine housing.

In this way, three such cooling segments are arranged side by side in the axial direction (FIG. 2). The mutual sealing can be done easily by inserting a sealing cord between two adjacent feet.

In the circumferential direction, the number of cooling segments 40 arranged next to one another corresponds to the number of front segments 31, so that a cooling segment is assigned to each front segment and to the burner 20 closest to the wall (FIG. 3). To form a closed cooling chamber 44, the cooling segments are also equipped with radially extending walls 45 in the circumferential direction. On the occasion of assembly, the cooling segments are brought into abutment with these walls 45. The end faces of the walls seal against the underside of the segment carrier 43

On its side facing away from the combustion chamber, i.e. on the side facing the cooling chamber 44, each cooling segment 40 is provided with a ribbed or corrugated surface 41. The ribs run in the circumferential direction (Fig. 2). In principle, the flow direction of the cooling air within the cooling chamber is thus predetermined.

A cooling segment is supplied with cooling air via a radially directed opening 46, which penetrates the segment carrier 43 and connects the collecting space 15 to one end of the cooling chamber 44 lying in the circumferential direction, as close as possible to the wall 45. Also located at the opposite end of this same cooling chamber the outlet opening 47 in the segment carrier as close as possible to the wall 45 there.

Both the opening 46 and the outlet opening 47 can either be individual bores or elongated holes that extend in the axial direction over a large part of the segment width.

The outlet opening 47 opens into a channel 48 which penetrates the segment carrier 43 in its entire axial extent and is open on both sides. On the turbine side, it opens against the annular space 35 between the double wall of the outer ring 34. As indicated schematically in FIG. 2, this outer ring is flanged to the segment carrier, the contour of the inner wall being matched to the contour of the cooling segments. On the burner side, the channel 48 opens against a head space 49, which is delimited by the cover 30 and the front segments 31. The cover 30 is also flanged to the segment carrier 43.

These axial channels 43, one of which is assigned to a segment in the circumferential direction, thus serve to jointly guide the segment cooling air and the cooling air acting on the secondary zone. Since the exit of the flame tube 33, 34 directed against the primary zone and the exit of the cooling segments accordingly lead directly into the combustion chamber inlet via the channels 48, the entire cooling air is fed to the combustion process without a large pressure drop.

The same measures are taken for cooling the inner wall of the primary zone as is indicated in FIG. 3 with the aid of the cooling segments 40 '.

Of course, the invention is not limited to the exemplary embodiment shown and described. It could just as well be used for wall cooling of pot-type combustion chambers.

Reference list

1

Gas turbine

2nd

compressor

3rd

Combustion chamber

10th

Machine axis

11

rotor

12th

Shovel carrier

13

Turbine casing

14

Exhaust housing

15

Gathering room

16

Inner part of the exhaust housing

17th

Outer part of the exhaust housing

18th

Flow rib

19th

Diffuser of 1

20th

burner

21

Support bearing

22

Diffuser of 2

23

drum

24th

Drum cover

25th

Ring channel

26

Partial body of 20

27

Partial body of 20

28

tangential slot

29

Fuel nozzle

30th

cover

31

Front segment

32

Secondary zone

33

Inner ring of 32

34

Outer ring of 32

35

Annulus of 34

36

Primary zone

37

Annulus of 33

38

Hub diffuser

39

Ring bowl

40, 40 '

Cooling segment

41

corrugated surface of 40

42

Foot

43

Segment carrier

44

Cooling chamber

45

radial wall of 40

46

opening

47

Outlet opening

48

channel

49

Headroom

Claims (6)

1. Gas-turbine combustion chamber having an annular combustion space (32, 36) whose walls extend from the combustion chamber inlet, whose circular-ring-shaped cross section is fitted with burners (20), to the inlet of the gas turbine (1) and are at the same time exposed to an air flow supplied by the compressor (2) of the gas turbine and are cooled by the latter, the cooling air being drawn at least partly from a collecting space (15) bounded by the turbine housing (13), a plurality of individually cooled cooling segments (40) forming the flow-limiting wall in a primary zone (36), and the cooling segments being lined up in the circumferential direction and in the axial direction and being suspended in a segment carrier (32),
   characterized in that

   - the combustion space is subdivided into the primary zone (36) and a secondary zone (32), whose flowlimiting walls (40, 33, 34) are separated and are cooled independently of one another,

   - in that the segment carrier (43) forms the outer boundary between the primary zone and the collecting space (15),

   - in that the secondary zone (32) situated downstream is bounded by a double-walled flame tube (33, 34) which surrounds an annular space (35, 37) and whose inlet end on the turbine side is open and forms the inlet for the cooling air of the secondary zone, the cooling air in the annular space (35, 37) flowing in the direction of the primary zone (36) in a counterflow with respect to the combustion chamber flow,

   - and in that both the outlet end, directed towards the primary zone (36), of the annular space (35, 37) of the flame tube (33, 34) and the outlet openings (47) of the cooling segments (40) of the primary zone are in conducting connection with the burners (20) at the combustion chamber inlet.
2. Gas-turbine combustion chamber according to Claim 1, characterized in that there are disposed in the segment carrier (43), on the one hand, radial openings (46), communicating with the collecting space (15), for the supply of the segment cooling air and, on the other hand, axial channels (48), communicating with the burner inlet, for the combined removal of the segment cooling air and of the cooling air acting on the secondary zone.
3. Gas-turbine combustion chamber according to Claim 2, characterized in that the number of channels (48) in the segment carrier (43) corresponds to the number of cooling segments (40) in the circumferential direction.
4. Gas-turbine combustion chamber according to Claim 1, in which the burners (20) are premix burners having two hollow conical subbodies (26, 27) which are nested in one another and whose central axes are offset with respect to one another such that in the longitudinal extension of the subbodies tangential slots (28) are formed, two burners in each case being radially disposed one on top of the other on a front segment (31) and a plurality of front segments lined up with one another forming a circular ring, characterized in that, in the circumferential direction, the number of cooling segments (40) lined up and the number of front segments (31) form an integral ratio.
5. Gas-turbine combustion chamber according to Claim 4, characterized in that, in the circumferential direction, the number of cooling segments (40) lined up corresponds to the number of front segments (31).
6. Gas-turbine combustion chamber according to Claim 2, characterized in that, in the axial direction, at least three cooling segments (40) disposed next to one another extend over the primary zone (36).

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