JPH06213458A - Combustor of gas turbine - Google Patents

Abstract

PURPOSE: To minimize consumption of cooling air to reduce the release of NOx . CONSTITUTION: A combustion chamber is divided into a primary zone and a secondary zone, and walls restricting flows in both the zones are cooled separately and in an unrelated fashion. A plurality of cooling segments 40 are individually cooled to form the walls restricting the flows and suspended on a segment supporting body 43. Moreover, the segment supporting body 43 forms a restricting part outside the primary zone with respect to a collection chamber 15 and the secondary zone positioned on the downstream side is restricted by a flame pipe having a double wall. In this manner, an inlet end part on the turbine side of the flame pipe is kept open to form an inlet for cooling air of the secondary zone. Thus, the outlet end part of the flame pipe directed to the primary zone and parts on the outlet side of the cooling segment 40 of the primary zone alone communicate with a burner 20 disposed at an inlet of a combustor.

Description

Detailed Description of the Invention

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a combustor of a gas turbine having an annular combustion chamber, the wall of the combustion chamber having a circular cross section with a burner and a gas turbine inlet to the gas turbine. Extending to the inlet of the gas turbine and surrounded by and cooled by an air stream supplied by a compressor of the gas turbine, the cooling air coming from a collecting chamber limited by the turbine casing. At least partially with respect to the type intended to be retrieved.

[0002]

Combustors for gas turbines with air-cooled flame tubes are known, for example from U.S. Pat. No. 4,077,205 or 3,978,662. The flame tube is mainly composed of wall portions overlapping in the turbine axis direction. On the side opposite to the combustion chamber, these wall parts each have a plurality of inlet openings distributed over the entire circumference. Through these inflow openings, air is introduced into the distribution chamber which is arranged in the flame tube and which is in communication with the combustion chamber. In this cooling system, each flame tube has a lip, which extends through a slit, through which a cooling air film emerges. The cooling film is adapted to adhere to the walls of the flame tube, thereby forming a cooling barrier layer for the flame tube.

Recently, "lean premixed combustion" is carried out for the combustion of gaseous fuel or liquid fuel with less harmful substances. In this case, the flame is supplied only when the fuel and the combustion air are premixed as uniformly as possible. When this is done with a high air excess, as is customary in gas turbine plants, a relatively low flame temperature results. This produces only the desired small amount of nitrogen oxides.

However, the above-described known gas turbine combustor has the following drawbacks. That is, the air consumption for cooling is extremely large, and further, the cooling air is supplied to the inside of the flame tube toward the downstream side of the flame, so that this air cannot be used for the original combustion process. Therefore, this combustor cannot operate with the high excess air factor required.

[0005]

OBJECTS OF THE INVENTION It is therefore an object of the present invention, in order to reduce the emission of NO _x, such as to minimize the cooling air consumption, a combustor in the form of a gas turbine mentioned at the outset Is to provide.

[0006]

In order to solve this problem, in the structure of the present invention, (a) the combustion chamber is divided into a primary zone and a secondary zone, and the primary zone and the secondary zone are The flow limiting walls are adapted to be cooled separately and independently of each other, and (b) in the primary zone, a plurality of individually cooled cooling segments form flow limiting walls. The cooling segment is suspended on a segment support, which forms a restriction outside the primary zone for the collection chamber, and (c) a secondary zone located downstream. Is limited by a flame tube having a double wall, the inlet end of the flame tube on the turbine side is open and forms the inlet for the cooling air in the secondary zone, and (d) the primary The flame facing the zone The outlet end of the tube and each portion of the primary zone on the outlet side of the cooling segment communicated only with the burner provided at the combustor inlet.

[0007]

The advantages of the invention are seen in particular in that the pressure drop of these cooling air streams can be kept small by the new measures based on the separate crossflow of both cooling air streams. Finally, this entire cooling air is fed to the combustion process after cooling has taken place. For introducing segment cooling air into the segment support, on the one hand,
A radial opening communicating with the collecting chamber is arranged, on the other hand, an axial passage communicating with the burner inlet for leading together the segment cooling air and the cooling air supplied to the secondary zone. The arrangement is particularly advantageous. In addition to this support function, the segment support also has the function of guiding the entire cooling air flow. Since this segment support is most often a cast member, the required openings can be manufactured very easily. Thereby, no additional air conduit is needed.

If a double-cone premixed burner is used as the burner, in most cases 2
Four burners are arranged radially above and below each other in the front segment. Burners of the front segments adjacent to each other are respectively offset in the radial direction for space reasons in these front segments formed in one annular body. As a result, the second burner in each circumferential direction is
It will be located close to the cooling segment. When viewed in the circumferential direction, the number of cooling segments arranged side by side is
Corresponding to the number of front segments, the number of air inflow openings and outflow passages in the segment support also corresponds to the number of circumferential cooling segments, for example by different sizing of the inflow or outflow holes. The amount of air introduced into the cooling segment can be easily adjusted according to the heat load of the cooling segment.

[0009]

Embodiments of the present invention will be described below with reference to the drawings showing a single-shaft axial flow gas turbine. Only the members important for understanding the present invention are shown. Of this device, for example, the complete state of the exhaust pipe with a flue, and
The inlet section of the compressor section is not shown. The flow direction of the operating medium is indicated by an arrow.

In FIG. 1, only the upper half of the machine axis 10 is shown in this apparatus. On the gas turbine 1 side, a rotor 11 in which rotating blades are arranged and a blade support in which guide blades are mounted are mainly mounted. It consists of body 12. The blade support 12 is inserted into a corresponding receiving portion provided in the turbine casing 13 via the protrusion and is suspended. An exhaust gas casing 14 is flange-connected to the turbine casing 13, and the exhaust gas casing mainly includes an annular inner portion 16 on the boss side and an annular outer portion 1.
7 and both parts are diffuser 19
Is restricted. These inner part 16 and outer part 1
7 is most often a half shell with an axial dividing surface. The two parts are connected to one another by a plurality of radial flow ribs 18, which are arranged evenly distributed over the entire circumference. A support on the outlet side of the turbomachine is arranged in the hollow chamber inside the inner portion 16, and the rotor 11 supports the support bearing 2
1 has been inserted.

Turbine casing 13 and blade support 12
Is provided with a horizontal dividing surface (not shown) located on the machine axis 10. Here, the upper and lower halves of the turbine casing and the vane support, which are often provided with flanges, are screwed together.

In the embodiment shown, the turbine casing 13 also encloses a collecting chamber 15 for the compressed combustion air. A part of the combustion air reaches the annular combustor 3 from the collecting chamber 15, and the combustor itself is opened to the turbine inflow portion, that is, to the upstream side of the first guide row. There is. Compressed air reaches the gathering chamber from the diffuser 22 of the compressor 2. Only the last three stages of this compressor are shown.
The compressor and turbine guide vanes are mounted on a common shaft formed as the rotor 11. The central axis of this axis is the longitudinal machine axis 10 of the gas turbine unit.
Is formed.

The shaft portion located between the turbine and the compressor is formed as a drum 23. This drum is
Its entire axial length is surrounded by a drum cover 24, which is fixed to the diffuser outer casing of the compressor via ribs (not shown). On the compressor side, this drum cover forms the cover band used for the blades of the last compressor guide row. On the turbine side, this drum cover, together with the end face of the turbine rotor, limits the radially extending impeller side chamber. This chamber forms the outlet end of the annular passage 25, which extends between the drum cover and the drum, starting from the boss behind the last compressor guide row. The entire cooling air on the rotor side is introduced into this annular passage.

The head edge of the combustor 3 is provided with a premixed burner 20, as is known, for example, from EP 321809. Such a premix burner (schematically shown in Figure 2) is a double cone burner. Mainly, this double cone burner
It consists of two hollow partial cones 26, 27 which are fitted in and out of one another in the direction of flow. In this case, the central axes of the two partial cones are offset from each other. Adjacent walls of these two partial cones form tangential slits 28 for the combustion air along the longitudinal length of the wall, so that this combustion air reaches the interior of the burner. It has become.
Here, a fuel nozzle 29 for liquid fuel is arranged. This fuel is injected into the hollow cone at a predetermined acute angle. The conical liquid fuel formation that occurs here is
It is surrounded by combustion air flowing in tangentially. In the axial direction, the fuel concentration is reduced by mixing with the combustion air. The burner may be operated with gaseous fuel. For this purpose, longitudinally distributed gas inlet openings are provided near the tangential slits in the walls of the two partial cones. As a result, during gas operation, the mixture formation with the combustion air already begins in the zone of the slit 28 for inflow. In this way, it can be seen that mixed operation with both fuel types is also possible. At the burner outlet, a very uniform fuel concentration is produced over the loaded annular cross section. At the burner exit,
A defined bulb-shaped backflow zone occurs, and ignition is performed at the tip of this backflow zone.

A part of the compressed combustion air is collected in the collecting chamber 15
Through the cover 30 with holes and flows into the burner in the direction of the arrow. During this combustion, the combustion gas reaches an extremely high temperature. This is a special requirement placed on the combustor wall to be cooled. Low NO _x
It is even more effective if a burner is used, for example the premix burner underlying this embodiment. These premix burners require a relatively small amount of cooling air. An annular combustion chamber extends downstream of the burner opening to the turbine inlet. The combustion chamber is bounded on the inside and outside by the wall to be cooled. These walls are often constructed as self-supporting structures.

In the above points, the annular combustor used in the gas turbine is known.

The combustor includes 72 burners 20. From FIG. 3, which shows a quadrant, the layout of this burner can be seen. Two burners are arranged on the front segment 31 one above the other in the radial direction. These 36 front segments side by side
It forms a closed annulus, which forms a heat shield. Both burners of the front segments adjacent to each other are offset in the radial direction. This can be seen from FIG.
This means that the radially outer burners on each second front segment are directly adjacent to the outer annular wall of the combustor. Therefore, the radially inner burner provided on the other front segment is arranged directly near the inner annular wall. From this,
A non-uniform heat load on the corresponding annular walls is created over the entire circumference.

The interior of the combustor is here divided into two zones, the walls of which are cooled differently from one another. A secondary zone 32 located downstream and opening to the turbine inlet is bounded by a double-walled flame tube. The inner ring 33 and the outer ring 34 of the flame tube consist of flangeless, welded sheet metal structures. The thin plate structure is closed by a spacer (not shown). Both ring bodies 3
3, 34 are open at the turbine end and form the inlet for the cooling air at this location.

As can be seen in FIG. 1, the outer annular body 34.
The annular chambers 35 between the double walls of the device draw air directly from the collecting chamber 15. Subject to efficient convection cooling, the air flows countercurrent to the combustor flow towards the primary zone 36.

Annular chamber 3 between the double walls of the inner annular body 33
Air is supplied to 7 from a boss diffuser 38.
The boss diffuser follows the diffuser 22 of the compressor and is limited by the drum cover 24 and the annular shell 39. The annular shell is connected to the drum cover 24 via ribs (not shown). Also in this annular chamber 37, the air flows countercurrent to the combustor flow towards the primary zone 36.

According to the invention, the cooling of the walls of the highly loaded primary zone is provided by the cooled individual cooling segments 40. These cooling segments, which are arranged side by side in the circumferential and axial directions, form the walls of the primary zone 36, which limit the flow over the entire axial length of the primary zone 36. Individual cooling has the advantage that there is little pressure drop. This cooling action
Can be adapted to local requirements.

This cooling segment 40 under high heat load
Is made of high heat resistant precision cast alloy. Each of these cooling segments 40 is provided with two bases 42 with supporting serrations, which bases 42 are circumferentially inserted and suspended in corresponding grooves provided in the support structure. ing. This is similar to, for example, the guide vane base being fixed to the vane support. Also similar to the vane support, this support structure (hereafter referred to as segment support 43) consists of two cast half shells with horizontal separation planes and pawls (not shown). Made of The support structure is supported within the turbine casing by the pawls.

In the axial direction, three such cooling segments are arranged next to one another (FIG. 2). Sealing of each other is facilitated by inserting a seal cord between two adjacent bases.

When viewed in the circumferential direction, the number of cooling segments 40 arranged side by side corresponds to the number of front segments 31, so that one cooling segment is provided for each front segment and the burner adjacent to the wall. Being assigned (Fig. 3). In order to form a closed cooling chamber 44, these cooling segments are provided with circumferentially, also radially extending walls 45. During assembly, these walls of the cooling segment abut each other. Each end surface of these walls seals toward the lower surface of the segment support 43.

On the side facing away from the combustion chamber, ie the side facing the cooling chamber 44, each cooling segment 40 is provided with a ribbed or corrugated surface 41. These ribs extend in the circumferential direction (FIG. 2). Thereby, in principle, the flow direction of the cooling air is defined inside the cooling chamber.

The supply of cooling air to the cooling segment takes place via radially oriented openings 46. This opening penetrates the segment support body 43, and
5 and the end of the cooling chamber 44 located in the circumferential direction are connected as close to the wall 45 as possible. At the opposite end of this same cooling chamber, an outlet opening 47 is arranged in the segment support, again as close as possible to the wall 45 at this location. Such openings 46 and outflow openings 47 are
It may be an individual hole or an elongated hole extending axially over most of the segment width.

The outflow opening 47 opens into the passage 48,
This passage runs through the segment support 43 along its entire axial length and is open on both sides. On the turbine side, this passage opens towards an annular chamber 35 provided between the double walls of the outer annular body 34. As shown diagrammatically in FIG. 2, this outer annulus is flanged to the segment support and the contour of the inner wall matches the contour of the cooling segment. On the burner side, the passage 48 opens towards the head chamber 49, which is
It is limited by the cover 30 and the front segment 31. The cover 30 is also flanged to the segment support 43.

One segment is assigned to each of the axial passages 48 in the circumferential direction. Thereby, these passages serve to guide together the cooling air of the segment and the cooling air supplied to the secondary zone. Thus, the outlets of the inner and outer annular bodies 33, 34, which are designed as flame tubes, towards the primary zone and the outlet of the cooling segment open directly via the passage 48 to the combustor inlet, The entire cooling air is fed to the combustion process without a large pressure drop.

To cool the inner wall of the primary zone, as indicated by cooling segment 40 'in FIG.
Similar means are used.

Of course, the invention is not limited to the embodiment described above. The combustor according to the invention can equally well be used for wall cooling of a pot-shaped combustor.

[Brief description of drawings]

FIG. 1 is a partial vertical cross-sectional view showing a gas turbine.

FIG. 2 is an enlarged view showing a portion of a primary zone of a combustor.

FIG. 3 is a partial cross-sectional view of the combustor primary zone taken along line 3-3 of FIG.

[Explanation of symbols]

1 gas turbine, 2 compressor, 3 combustor, 1
0 machine axis, 11 rotor, 12 blade support,
13 turbine casing, 14 exhaust gas casing, 15 collecting chamber, 16 inner part, 17 outer part, 18 flow rib, 19 diffuser, 20
Burner, 21 support bearing, 22 diffuser, 23 drum, 24 drum cover, 25
Annular passage, 26, 27 partial cone, 28 slit, 29 fuel nozzle, 30 cover, 31 front segment, 32 secondary zone, 33, 34
Annular body, 35 annular chamber, 36 primary zone, 37
Annular chamber, 38 boss diffuser, 39 annular shell, 40, 40 'cooling segment, 41 corrugated surface, 42 base, 43 segment support, 44
Cooling chamber, 45 walls, 46 openings, 47 outflow openings, 48 passages, 49 head chambers

Claims (6)

[Claims] 1. A combustor of a gas turbine with an annular combustion chamber (32, 36), the wall of the combustion chamber being from an inlet of an annular cross section with a burner (20), It extends to the inlet of the gas turbine (1), is surrounded by and is cooled by the air stream supplied from the compressor (2) of the gas turbine, the cooling air being the turbine. Of the type adapted to be at least partially taken out of the collecting chamber (15) limited by the casing (3), (a) the combustion chamber comprises a primary zone (36) and a secondary zone (3).
2) and the flow-limiting walls (40, 33, 34) of the primary and secondary zones are cooled independently of each other and independently of each other, B) In the primary zone (36), a plurality of individually cooled cooling segments (40) form flow limiting walls, which cooling segments (4).
3) is inserted and suspended, the segment support forms a restriction portion outside the primary zone for the collecting chamber (15), and (c) the secondary zone (32) located on the downstream side. ) Is bounded by a double-walled flame tube (33, 34), the turbine-side inlet end of which is open to form an inlet for the secondary zone cooling air. (D) The flame tube (3) facing the primary zone (36)
3, 34) and the outlet-side parts of the cooling zone (40) of the primary zone are in communication only with a burner (20) provided at the combustor inlet, Gas turbine combustor. 2. Collecting chamber (15) for introducing segment cooling air into the segment support (43), on the one hand.
A radial opening (46) communicating with
On the other hand, an axial passage (48) communicating with the burner unit for deriving together the segment cooling air and the cooling air supplied to the secondary zone is arranged,
The combustor according to claim 1. 3. Combustor according to claim 2, wherein the number of passages (48) provided in the segment support (43) corresponds circumferentially to the number of cooling segments (40). 4. The burner (20) is a premixed burner of double-cone structure, wherein two burners each are arranged on each front segment (31) in a radial direction one above the other and adjacent to one another. Of the front segments forming one annular body, the number of the cooling segments (40) arranged side by side in the circumferential direction and the number of the front segments (31) form a predetermined integer ratio. The combustor of claim 1, wherein: 5. Combustor according to claim 4, wherein the number of cooling segments (40) arranged side by side in the circumferential direction corresponds to the number of front segments (31). 6. The combustor according to claim 2, wherein at least three cooling segments (40) arranged axially adjacent to each other extend over the primary zone (36).