WO2013056819A1

WO2013056819A1 - Lean premix burner of an aircraft gas turbine engine

Info

Publication number: WO2013056819A1
Application number: PCT/EP2012/004327
Authority: WO
Inventors: Imon-Kalyan Bagchi; Waldemar Lazik
Original assignee: Rolls-Royce Deutschland Ltd. & Co Kg
Priority date: 2011-10-18
Filing date: 2012-10-16
Publication date: 2013-04-25
Also published as: DE102011116317A1

Abstract

The invention relates to a lean premix burner of an aircraft gas turbine engine, comprising a centrally arranged support burner (29) and a main burner (30) that annularly surrounds the support burner, characterized in that the support burner (29) has a primary fuel channel (31) and a secondary fuel channel (32), which fuel channels can be supplied with fuel independently of each other.

Description

Magervormian burner of an aircraft gas turbine engine

description

The invention relates to a lean burn burner of an aircraft gas turbine engine with a centrally arranged support burner and a torch surrounding this annular main burner.

It is known to use two fuel atomizers in lean burn burners, namely a support burner and a main burner. The support burner is arranged centrally in the main burner. The support burner is usually designed as Druckdrallzerstäuber. The lean burn burner includes two fuel lines to supply the support burner and the main burner. In operation, the backup burner is used to ignite the aircraft gas turbine engine and at low load conditions, while the main burner is operated at partial load and is used up to maximum load. The support burner is designed for the ignition process and for a stable combustion during the starting phase of the engine. This requires a relatively small flow rate in the case of the pressure swirl atomizer. This is in contrast to the need to use the support burner for a larger operating range until the transition to large fuel mass flows. From EP 2 051 010 A1 a lean premix burner is known, which has a conical film applicator with a smooth surface. The film applicator is supplied by a fuel supply device which forms a continuous film by twisting the fuel. The fuel is also applied to the outside of the film layer. The associated air flow on the fuel side has the same spin as the fuel. The cross sections of the annular air channels decrease to ensure acceleration of the fuel. Another lean burn burner is previously known from EP 1 801 504 A2. This lean burn burner is internally graded. The main burner is an airflow filer, the surface of which has a plurality of channels arranged in the flow direction to produce a plurality of narrow fuel streams to ensure homogeneity in the circumferential direction.

The invention has for its object to provide a lean burn burner of the type mentioned, which ensures optimum combustion with a simple design and safe operation over a large power range of the aircraft gas turbine engine.

According to the invention the object is achieved by the combination of features of claim 1, the dependent claims show further advantageous embodiments of the invention. According to the invention it is thus provided that the support burner has a central primary fuel channel and a surrounding annular secondary fuel channel, which are supplied with fuel independently of each other. In contrast to the prior art drawbacks and overcoming the same, a lean burn burner has thus been provided in accordance with the present invention which ensures stable combustion of the fuel over a wide performance range of the aircraft gas turbine engine. To ensure good ignition of the combustion chamber, the support burner must be able to optimally atomize the fuel at low fuel mass flows (flow rates). This requires a separate fuel supply of the support burner with small amounts of fuel. According to the invention this is done by the centric primary fuel channel, which is connected to a fuel supply system (fuel pump, distribution channels and the like). In contrast, in the constructions known from the prior art, it is necessary to ignite the main burner already at very low powers of the gas turbine engine, which may result in an insufficient fuel supply in the prior art, which can lead to unfavorable emission levels of the main burner. This disadvantage has been overcome according to the invention in that different fuel supply systems are used, namely via the primary fuel channel and the secondary fuel channel.

According to the invention, a further disadvantage of the prior art is avoided in that a good ignition of the support burner can take place, while in the constructions known from the prior art can result in disadvantages in the ignition, since the support burner must be designed for a wide operating range. In the known designs, the fuel pressure in the fuel system must be limited in order to operate the support burner in the start-up phase of the aircraft gas turbine engine. A prior art support burner with larger fuel mass flows results in poor atomization in the ignition condition and at low load conditions. Therefore, in the systems known from the prior art compromises are required that partially preclude optimal operation. This situation is overcome by the solution according to the invention.

Furthermore, it is possible according to the invention not only to atomize the fuel in a suitable manner, but also to initiate optimally in the combustion chamber. This results in considerably improved emission values according to the invention compared with the prior art. According to the invention thus a total of three fuel circuits are provided, namely two fuel circuits for the support burner (primary fuel channel and secondary fuel channel) and a fuel circuit for the main burner. The primary fuel system is used for the ignition and low load conditions of the aircraft gas turbine engine, while the secondary fuel system is used at higher load conditions. The activation of the two fuel systems can be done according to the invention either actively, for example by electrical control or passively, for example by fuel pressure-dependent valves. According to the invention results from the dual design of the support burner improved atomization and from this improved efficiency throughout the operating range. The improved control of the combustion stoichiometry in the backup combustor zone results in a further reduction in exhaust emissions.

While the primary fuel supply contributes to good atomization at low fuel flow rates and good ignition, the secondary fuel supply is switched on at higher operating conditions (greater power of the aircraft gas turbine engine). The secondary fuel system has a higher flow rate. As a result, the possibility is created to operate the support burner even at higher operating conditions of the aircraft gas turbine engine, without the fuel pressure in the fuel system exceeds the predetermined limits. The larger mass flow in the secondary fuel system does not reduce the quality of the atomization of the fuel since the fuel mass flow is already relatively high when the secondary fuel system is switched on. The air temperature and pressure are also high enough to ensure good secondary atomization.

In the following the invention will be described by means of embodiments in conjunction with the drawing. Showing:

1 is a schematic representation of a gas turbine engine according to the present invention, Fig. 2 is a schematic representation of the invention

Magervormischbrenners,

Fig. 3 is a schematic representation of the combustion zones in the

Combustion chamber with the lean burn burner shown in FIG. 2,

4 shows a detailed representation of a first embodiment of the lean burn burner according to the invention, FIG. 5 shows a modified embodiment of that shown in FIG. 4

Embodiment, another embodiment of the lean burn burner according to the invention, another embodiment of the lean burn burner according to the invention, another embodiment of the lean burn burner according to the invention,

9 is a detail view of FIG. 8,

Fig. 10 is a sectional view taken along the line A-A of Fig. 9 and

Fig. 1 1, 12 representations of possible operations of the support burner and the

Main burner.

The gas turbine engine 10 of FIG. 1 is an example of a turbomachine to which the invention may apply. However, it will be understood from the following that the invention can be used with other turbomachinery as well. The engine 10 is formed in a conventional manner and comprises in the flow direction one behind the other an air inlet 1 1, a circulating in a housing fan 12, an intermediate pressure compressor 13, a high pressure compressor 14, combustion chambers 15, a high-pressure turbine 16, an intermediate-pressure turbine 17 and a low-pressure turbine 18 and an exhaust nozzle 19, all of which are arranged about a central engine axis 1. The intermediate pressure compressor 13 and the high pressure compressor 14 each include a plurality of stages, each of which includes a circumferentially extending array of fixed stationary vanes 20, commonly referred to as stator vanes, that radially inwardly from the engine casing 21 in an annular flow passage through the compressors 13, 14 protrude. The compressors further include an array of compressor blades 22 projecting radially outwardly from a rotatable drum or disc 26 coupled to hubs 27 of the high pressure turbine 16 and the intermediate pressure turbine 17, respectively.

The turbine sections 16, 17, 18 have similar stages, comprising an array of fixed vanes 23 projecting radially inward from the housing 21 into the annular flow passage through the turbines 16, 17, 18, and a downstream array of turbine blades 24 projecting outwardly from a rotatable hub 27. The compressor drum or compressor disk 26 and the vanes 22 disposed thereon and the turbine rotor hub 27 and the turbine blades 24 disposed thereon rotate about the engine axis 1 during operation. The reference numeral 28 shows an outlet cone.

2 shows the basic structure of a lean burn burner according to the invention. This comprises a centric supporting burner 29 and an annular main burner 30 surrounding it. Centrally around the supporting burner 29 extends an annular air duct 33, which is bounded radially outwards by a flame stabilizer 34. Radially outside an annular air passage 35 is provided. In the air channels 33 and 35 each swirl generator 36 and 37 are arranged.

Concentric with the arrangement described, as mentioned, the main burner 30 is arranged. This is surrounded by an annular air channel 38 with a swirl generator 39. The reference numeral 40 shows an outer surface of the flame stabilizer 34. The main burner 30 includes a film layer 41 having a Ring channel 42 for introducing fuel. The air channel 38 is bounded by a bevelled ring 43.

The air for the secondary fuel of the support burner is passed through the annular air passage 33, while the air for the main burner is passed through the annular air passage 35.

The fuel system for the main burner is designed as a flow filer. The fuel is atomized by the air supplied through the two annular air passages 35 and 38. The supplied swirl generators 37 and 39 allow the supplied air (air passages 35 and 38) to rotate either in the same direction or in the opposite direction. The fuel exits through the annular channel 42, the total flow then follows the outer film layer 41st

Fig. 2 shows an annular flame stabilizer 34 for the support zone, which has a V-shaped cross-sectional area. The outer surface 40 of the flame stabilizer 34 supports flow along the fuel side of the main burner fuel atomizer. The configuration and the angle of the outer surface 40 are chosen so that there is a constant reduction of the flow region along the air channel 35. The radially outermost component is formed by the ring 43 with the chamfer. The ring 43 supports the entire assembly and forms the transition to the combustion chamber 15. The inner contour of the tapered ring 43 is formed so that it ensures a good flow of air on the outside of the film layer. As a result, the ignition properties of the burner are improved.

FIG. 3 shows the different flow fields which result in the construction shown in FIG. In this case, depending on the operating conditions, combustion occurs in the zones described below: A support flow 44 forms a jet which branches radially outwards. A main flow 46 results in the formation of a centric recirculation flow 47 in which the support flow is trapped. The recirculation flow 47 leads to a backflow against the flow direction to a stagnation point 45 at which the support flow 44 breaks off. This creates a stagnation zone at the stagnation point 45.

Furthermore, a flame stabilization zone 48 is formed as a result of the flow around the V-shaped flame stabilizer 34. The support flow 44 moves radially outward. The counter-swirl formed by the inner main swirler 36 results in a toroidal recirculation zone as indicated by reference numeral 48. The main flow 46 is formed by the flows through the air passages 33 and 35.

According to the invention, it is possible to realize the dual design of the support burner by different constructions. It is possible to provide two concentric Druckdrallzerstäuber. It is also possible to provide a pressure swirl atomizer for the primary support function and a compressed air atomizer for the secondary support function. Alternatively, two concentric Druckdrallzerstäuber can be provided with a common outflow. Another variant consists of a primary pressure swirl atomizer and a secondary single jet atomizer. These different design variants will be described below.

4 shows an embodiment with two concentric pressure swirl atomizers. The support burner 29 has a primary fuel channel 31, which is arranged centrally, and an annular secondary fuel channel 32. In the primary fuel passage 31, an array of annular slots or recesses are provided as fuel swirlers 49. The primary fuel passage 31 branches as shown in FIG. 4 and flows as a hollow cone jet through a circular outlet 50. The metering or point of greatest pressure drop can be provided either at the swirl generator 49 or at the outlet 50. The number of fuel swirler slots or recesses may be between three and eight, depending on the desired flow rate. The secondary fuel is swirled by an array of annular slots or recesses 51. The number of slots or recesses may also be between three and eight. The secondary fuel exits through an annular outlet 52. The fuel will atomized and mixed with the swirling assist air flow passing through the axial swirl generator 36. A portion of the support flame stabilizes downstream of the flame stabilizer 34. The secondary fuel flow outlet may be formed as shown in FIG. 5 to permit selective separation of the fuel output between the primary and secondary fuel flows. This improves the control of the local stoichiometry. Fig. 6 shows an embodiment with a Druckdrallzerstäuber as a primary support function and a Druckluftzerstäuber as a secondary support function. The fuel is supplied through the central primary fuel passage 31 and the secondary annular fuel passage 32. The primary fuel is swirled by an array of annular slots or recesses (fuel delivery system 49). The number of slots or recesses may be between three and eight, depending on the desired fuel mass flow. The primary fuel exits as a hollow cone jet through the outlet 50. The metering or maximum pressure loss point may be at either the fuel rail 49 or outlet 50.

The secondary atomizer is designed as a compressed air atomizer. The secondary fuel is twisted by an array of annular slits or recesses (fuel rail producer 51). The number of swirler slots or recesses may be between three and eight, depending on the desired mass flow. The secondary fuel exits through an annular outlet 52. The secondary atomizer is designed as a compressed air atomizer and dimensioned such that a radially inner surface 53 forms the primary air channel 33. The primary air can be supplied without swirl or using an axial swirl generator. The secondary fuel is atomized and mixed with the two annular air flows, namely the radially inner air flow (air passage 33) and the radially outer air flow (air passage 54). A part of the support flame stabilizes downstream of the flame stabilizer 34. In the embodiment shown in FIG. 7, two concentric pressure swirl atomizers with a common outlet 50 are provided. The fuel is supplied through the central primary fuel passage 31 and through the annular secondary fuel passage 32. The primary fuel is swirled using an array of annular slots or holes (fuel swirler 49). The number of spin-producing slots or holes may be between three and eight, depending on the desired fuel mass. The fuel exits as a hollow cone jet through the circular outlet 50. The metering or maximum pressure drop point is at the fuel swirler 49 and not at the outlet 50. The flow area of the outlet 50 is sized to be the metering or throttling point as the primary fuel stream and the secondary fuel stream flow out.

The secondary atomizer is a pressure swirl atomizer. The secondary fuel is swirled by an array of annular slots or holes (fuel swirler 55). The number of spin slots or holes may be between three and eight, depending on the desired fuel mass. The secondary fuel also exits through the circular outlet 50 as a hollow cone jet. Both jets of fuel are atomized and mixed with the swirl air flow supplied by the axial swirl generator 36. A part of the support flame stabilizes downstream of the flame stabilizer 34.

Fig. 8 shows another embodiment in which a primary pressure swirl atomizer and a secondary single-jet atomizer are used. The primary fuel is spun using an array of annular slits or swirl recesses, the number of which is between three and eight, depending on the desired mass flow of the fuel. The primary fuel forms a hollow cone spray through the circular outlet 50. The metering or point of highest pressure drop is at either the fuel swirler 49 or the outlet 50. The secondary fuel is supplied as a single jet injection through an array of recesses 56. These recesses can either be arranged only radially, be formed with tangential orientation with axial orientation. The number of recesses can vary between three and eight. The primary fuel is radially outside the Secondary fuel arranged to cool the individual beam recesses 56. Both fuel streams are atomized and mixed with the twisted primary air flow supplied by the axial swirl generator 36. This is shown in detail in FIG. 9. FIG. 10 shows an enlarged sectional view along the line AA of FIG. 9.

In the embodiment variants described in connection with FIGS. 6 and 7, there results an improved interaction between the support burner air and the primary fuel. Because there is no concentric secondary fuel flow, the air can reach the primary fuel at the fuel exit, which improves secondary atomization and mixture formation.

In the embodiment described in connection with FIG. 6, the secondary region of the support burner has the advantages of a compressed air atomizer, in particular with regard to the formation of soot. In the exemplary embodiment shown in FIGS. 8 to 10, the individual, discrete beams lead to a good atomization quality. The radially outwardly directed individual rays lead to a higher beam penetration with increasing fuel flow. Thus, as the load of the gas turbine engine increases, the fuel assembly is moved radially outward. This leads to an improvement of the mixture formation.

FIGS. 1 and 12 show possible procedures for the operation of the two-stage support burner 29 according to the invention and of the main burner 30. In each case, the proportion of fuel over the engine load is shown.

In the systems known from the prior art with a single single-stage auxiliary burner, this can only be operated below an average power of the gas turbine engine. If this range is exceeded, high emissions result, in particular high soot development by the auxiliary burner and unfavorable combustion efficiency. Turning on the main burner at these low load conditions results in poor combustion efficiency of the main burner. As described, it has been proposed overcome this problem by a stepped design of the main burner, but this can significantly affect the performance of the turbine.

In the two procedures described below, the subdivision of the backup burner enables the supply of a primary fuel channel and a secondary fuel channel, so that the auxiliary burner can be operated at higher load conditions (60% and higher). Good ignition characteristics and starting characteristics of the gas turbine are ensured by a rich fuel supply in the first stage of the auxiliary burner (fuel supply through the primary fuel channel). Leaner combustion through the secondary stage of the backup burner (fuel supply through the secondary fuel channel) results in improved emissions performance at higher load conditions of the gas turbine engine. As a result, the main burner can be switched on only when a higher load condition is reached, whereby its efficiency is improved.

FIG. 11 shows a first option in which the first stage of the support burner 29 (supplied by the primary fuel passage 31) is used for ignition and start up to about 20% load of the gas engine. This leads to a fat stable flame. In the next stage, both areas of the support burner 29 are used (fuel supply through the primary fuel passage 31 and the secondary fuel passage 32). This level ranges between about 20% and about 60% load. The result is a rich flame from the first stage of the support burner (primary fuel channel) and a lean flame from the second stage of the support burner (secondary fuel channel). Between about 60% load and 100% load, the first stage of the support burner 29 and the main burner 30 are put into operation.

The variant described in FIG. 12 shows in the third area the additional use of the second stage of the supporting burner (secondary fuel channel) together with the main burner.

1 and 12 are as follows: Pilot 1 refers to the first stage of the support burner 29 with fuel supply through the primary fuel passage 31, Pilot 2 refers to the second stage of the support burner 29 Fuel supply through the secondary fuel channel, the main stage refers to the main burner.

LIST OF REFERENCES:

1 engine axis

10 gas turbine engine

1 1 air intake

12 circulating fan in the housing

13 intermediate pressure compressor

14 high pressure compressor

15 combustion chambers

16 high-pressure turbine

17 intermediate pressure turbine

18 low-pressure turbine

19 exhaust nozzle

20 vanes

21 engine case

22 compressor blades

23 vanes

24 turbine blades

26 Compressor drum or disc

27 turbine rotor hub

28 outlet cone

29 support burner

30 main burners

31 primary fuel channel

32 secondary fuel channel

33 air duct

34 flame stabilizer

35 Annular air duct

36 swirl generator

37 swirl generator

38 air duct

39 swirl generator

40 Outer surface of the flame stabilizer

41 film-makers annular channel

ring

support flow

stagnation point

mainstream

Recirculation flow flame stabilization zone fuel swirl generator outlet

Fuel swirl generation outlet

surface

air duct

Fuel swirler recess

Claims

claims

1 . Magervormischbrenner an aircraft gas turbine engine with a centrally arranged support burner (29) and a surrounding annular main burner (30), characterized in that the support burner (29) has a Primärbrennstoffkanal (31) and a Sekundärbrennstoffkanal (32), which are independently supplied with fuel ,

2. lean burn burner according to claim 1, characterized in that the central primary fuel passage (31) is designed for a fuel gas flow rated for ignition of the aircraft gas turbine engine.

3. lean burn burner according to claim 1 or 2, characterized in that the annular secondary fuel passage (32) is designed for a, for the part-load operation of the aircraft gas turbine engine, rated fuel mass flow.

4. lean burn burner according to one of claims 1 to 3, characterized in that the support burner (29) comprises two concentric swirl-associated pressure atomizer.

5. lean burn burner according to claim 4, characterized in that the atomizers comprise a common outlet region (50).

6. lean burn burner according to one of claims 1 to 3, characterized in that the support burner (29) comprises a with the primary fuel passage (31) connected to the centric swirl atomizer and connected to the secondary fuel passage (32) Druckluftzerstäuber.

7. lean burn burner according to one of claims 1 to 3, characterized in that the support burner (29) comprises a central, with the primary fuel channel (31) connected to the pressure atomizer and a surrounding, with the secondary fuel channel (32) connected to the single jet atomizer.

8. lean burn burner according to one of claims 1 to 7, characterized in that the primary fuel passage (31) is arranged centrally and by an annular secondary fuel passage (32) is surrounded.

9. lean burn burner according to one of claims 1 to 8, characterized in that the secondary fuel channel (32) is arranged centrally and surrounded by an annular primary fuel channel (31).

10. A method for operating a lean premix burner according to one of claims 1 to 9, wherein for firing and starting the gas turbine engine, a first stage of the support burner (29) with fuel supply through the primary fuel channel (31) up to a load of the gas turbine of about 20% in addition, between a load of about 20% to about 60% of the gas turbine, a second stage of the backup burner (29) is used with fuel supply through the secondary fuel channel (32) and between about 60% to about 60% load 100% of the gas turbine engine, the first stage of the support burner (29) and the main burner or the first stage and the second stage of the support burner and the main burner are used.