DE102011116317A1 - Magervormian burner of an aircraft gas turbine engine - Google Patents

Abstract

The invention relates to a lean burn burner of an aircraft gas turbine engine with a centrically arranged support burner 29 and a main burner surrounding this annular main burner 30, characterized in that the support burner 29 has a primary fuel passage 31 and a secondary fuel passage 32, which can be supplied with fuel independently.

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 the EP 2 051 010 A1 For example, a lean premix burner is known which has a conical sizer 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 from the EP 1 801 504 A2 previously known. 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 disadvantages resulting from the prior art and overcoming the same, according to the invention, therefore, a lean premix burner has been provided which ensures stable combustion of the fuel over a wide power range of the aircraft gas turbine engine. To ensure good ignition of the combustion chamber, the support burner must be in 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, the fuel not only in a suitable manner atomise, 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 according to the invention either active z. B. by electrical control or passive z. B. 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 performance 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 a schematic representation of a gas turbine engine according to the present invention,

2 a schematic representation of the lean burn burner according to the invention,

3 a schematic representation of the combustion zones in the combustion chamber with the in 2 Magervormischbrenner shown,

4 a detailed representation of a first embodiment of the lean burn burner according to the invention,

5 a modified embodiment of the in 4 shown embodiment,

6 a further embodiment of the lean burn burner according to the invention,

7 a further embodiment of the lean burn burner according to the invention,

8th a further embodiment of the lean burn burner according to the invention,

9 a detailed view according 8th .

10 a sectional view taken along the line AA of 9 and

11 . 12 Representations of possible operations of the support burner and the main burner.

The gas turbine engine 10 according to 1 is an example of a turbomachine to which the invention may find application. 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 one air inlet in succession in the flow direction 11 , 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 around a central engine axis 1 are arranged.

The intermediate pressure compressor 13 and the high pressure compressor 14 each comprise a plurality of stages, each of which comprises a circumferentially extending array of fixed stationary vanes 20 commonly referred to as stator blades and radially inward of the engine casing 21 in an annular flow channel through the compressors 13 . 14 protrude. The compressors further include an array of compressor blades 22 on, which is radially outward from a rotatable drum or disc 26 project with hubs 27 the high-pressure turbine 16 or the intermediate-pressure turbine 17 are coupled.

The turbine sections 16 . 17 . 18 have similar steps, comprising an array of fixed vanes 23 extending radially inward from the housing 21 into the annular flow channel through the turbines 16 . 17 . 18 project, and a subsequent arrangement of turbine blades 24 facing outward from a rotatable hub 27 protrude. The compressor drum or compressor disk 26 and the blades arranged thereon 22 as well as the turbine rotor hub 27 and the turbine blades disposed thereon 24 rotate around the engine axis during operation 1 , The reference number 28 shows an outlet cone.

The 2 shows the basic structure of a lean burn burner according to the invention. This includes a centric support burner 29 and a surrounding annular main burner 30 , Centric around the support burner 29 extends an annular air duct 33 , which radially outward through a flame stabilizer 34 is limited. Radially outside is an annular air duct 35 intended. In the air channels 33 and 35 are each swirl generator 36 respectively. 37 arranged.

Concentric with the arrangement described is, as mentioned, the main burner 30 arranged. This is from an annular air duct 38 with a swirl generator 39 surround. The reference number 40 shows an outer surface of the flame stabilizer 34 , The main burner 30 includes a filmmaker 41 with a ring channel 42 for the introduction of fuel. The air duct 38 is through a bevelled ring 43 limited.

The air for the secondary fuel of the support burner is through the annular air duct 33 passed through, while the air for the main burner through the annular air duct 35 to be led.

The fuel system for the main burner is designed as a flow filer. The fuel is atomized by the air passing through the two annular air channels 35 and 38 is supplied. Through the provided swirl generator 37 and 39 can the supplied air (air ducts 35 and 38 ) rotate either in the same direction or in the opposite direction. The fuel passes through the annular channel 42 out, the total flow then follows the outer film layer 41 ,

The 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 a flow along the fuel side of the main burner fuel atomizer. The design and the angle of the outer surface 40 are chosen so that there is a constant reduction of the flow area along the air channel 35 results. The radially outermost component is through the ring 43 formed with the bevel. The ring 43 stores the entire assembly and forms the transition to the combustion chamber 15 , The inner contour of the bevelled ring 43 is designed to ensure good air flow on the outside of the film layer. As a result, the ignition properties of the burner are improved.

The 3 shows the different flow fields, which at the in 2 shown construction. Depending on the operating conditions, combustion occurs in the zones described below:
A supporting flow 44 forms a beam which branches radially outwards. A mainstream 46 leads to the formation of a centric recirculation flow 47 in which the supporting flow is enclosed. The recirculation flow 47 leads to a backflow against the flow direction to a stagnation point 45 at which the support flow 44 aborts. This creates a stagnation zone at the stagnation point 45 ,

Furthermore, a flame stabilization zone is formed 48 as a result of the flow around the V-shaped flame stabilizer 34 , The supporting flow 44 moves radially outwards. By the counter-twist, which by the inner main swirl generator 36 is formed, resulting in a toroidal recirculation zone, as by the reference numeral 48 shown. The mainstream 46 is due to the currents through the air channels 33 and 35 educated.

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.

The 4 shows an embodiment with two concentric Druckdrallzerstäubern. The support burner 29 has a primary fuel channel 31 , which is arranged centrally, and an annular secondary fuel channel 32 on. In the primary fuel channel 31 are an array of annular slots or recesses as a fuel swirler 49 intended. The primary fuel channel 31 branches, as in 4 represented, and flows through as a hollow cone beam a circular outlet 50 , The metering or point of greatest pressure drop can be either at the swirl generator 49 or at the outlet 50 be provided. The number of fuel swirler slots or recesses may be between three and eight, depending on the desired flow rate. The secondary fuel is through an array of annular slots or recesses 51 The number of slots or recesses may also be between three and eight. The secondary fuel passes through an annular outlet 52 out. The fuel is atomized and mixed with the swirling support airflow passing through the axial swirl generator 36 flows. Part of the support flame stabilizes downstream of the flame stabilizer 34 ,

The outlet of the secondary fuel flow may be as in 5 shown to be formed to allow a targeted separation of the fuel output between the primary and secondary fuel flows. This improves the control of the local stoichiometry.

The 6 shows an embodiment with a Druckdrallzerstäuber as a primary support function and a compressed air atomizer as a secondary support function. The fuel is through the centric primary fuel channel 31 and the annular secondary fuel channel 32 fed. The primary fuel is passed through an array of annular slots or recesses (fuel swirl generators 49 ) subjected to spin. The number of slots or recesses may be between three and eight, depending on the desired fuel mass flow. The primary fuel passes through the outlet as a hollow cone jet 50 out. The metering or maximum pressure drop point may be at either the fuel swirler 49 or at the outlet 50 available.

The secondary atomizer is designed as a compressed air atomizer. The secondary fuel is twisted by an array of annular slots or recesses (fuel swirl generator 51 ). The number of swirler slots or recesses may be between three and eight, depending on the desired mass flow. The secondary fuel passed through an annular outlet 52 out. The secondary atomizer is designed as a compressed air atomizer and dimensioned such that a radially inner surface 53 the primary air channel 33 forms. 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 duct 33 ) and the radially outer air flow (air duct 54 ). Part of the support flame stabilizes downstream of the flame stabilizer 34 ,

At the in 7 shown embodiment are two concentric Druckdrallzerstäuber with a common outlet 50 intended. The fuel is through the central primary fuel channel 31 as well as through the annular secondary fuel channel 32 fed. The primary fuel is swirled using an array of annular slots or holes (fuel swirl generator 49 ). The number of spin-producing slots or holes may be between three and eight, depending on the desired fuel mass. The fuel passes as a hollow cone jet through the circular outlet 50 out. The metering or point of highest pressure drop is at the fuel swirl generator 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 swirl generator 55 ). The number of spin slots or holes may be between three and eight, depending on the desired fuel mass. The secondary fuel also passes through the circular outlet 50 as a hollow cone beam. Both jets of fuel are atomized and mixed with the swirl air flow, which through the axial swirl generator 36 is supplied. Part of the support flame stabilizes downstream of the flame stabilizer 34 ,

The 8th 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 either at the fuel swirl generator 49 or the outlet 50 , The secondary fuel is injected as a single jet through an array of recesses 56 fed. 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 disposed radially outward of the secondary fuel around the individual jet recesses 56 to cool. Both fuel streams are atomized and mixed with the twisted primary air flow passing through the axial swirl generator 36 is supplied. This is in the detail 9 shown. The 10 shows an enlarged sectional view taken along the line AA of 9 ,

In connection with the 6 and 7 described design variants results in an improved interaction between the supporting 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.

When in connection with 6 described embodiment, the secondary region of the support burner on the benefits of a Druckluftzerstäubers, in particular with regard to the formation of soot. In the in the 8th to 10 In the embodiment shown, 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.

The 11 and 12 show possible procedures of the operation of the two-stage support burner according to the invention 29 and the main burner 30 , In each case, the fuel fraction is shown above the engine load.

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 to overcome this problem by means of a stepped configuration of the main burner, which, however, can considerably impair the operating behavior 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.

The 11 FIG. 12 shows a first option wherein the ignition and start up to about 20% load of the gas engine is the first stage of the backup burner 29 (Supply through the primary fuel channel 31 ) is used. This leads to a fat stable flame. In the next stage, both areas of the support burner 29 used (fuel supply through the primary fuel channel 31 and the secondary fuel channel 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 become the first stage of the support burner 29 as well as the main burner 30 put into operation.

In the 12 described variant shows in the third area, the additional use of the second stage of the support burner (secondary fuel channel) together with the main burner.

The labels in the 11 and 12 are as follows: Pilot 1 refers to the first stage of the support burner 29 with fuel supply through the primary fuel channel 31 , Pilot 2 refers to the second stage of the support burner 29 with fuel supply through the secondary fuel channel, the main stage refers to the main burner.

LIST OF REFERENCE NUMBERS

1

Engine axis

10

Gas turbine engine

11

air intake

12

in the housing revolving fan

13

Intermediate pressure compressor

14

High pressure compressor

15

combustors

16

High-pressure turbine

17

Intermediate pressure turbine

18

Low-pressure turbine

19

exhaust nozzle

20

vanes

21

Engine casing

22

Compressor blades

23

vanes

24

turbine blades

26

Compressor drum or disc

27

Turbinenrotornabe

28

outlet cone

29

supporting burner

30

main burner

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 applicator

42

annular channel

43

ring

44

support flow

45

stagnation point

46

mainstream

47

recirculation

48

Flame stabilization zone

49

Fuel swirl generator

50

outlet

51

Fuel swirl generation

52

outlet

53

surface

54

air duct

55

Fuel swirl generator

56

recess

QUOTES INCLUDE IN THE DESCRIPTION

This list of the documents listed by the applicant has been generated automatically and is included solely for the better information of the reader. The list is not part of the German patent or utility model application. The DPMA assumes no liability for any errors or omissions.

Cited patent literature

• EP 2051010 A1 [0003]
• EP 1801504 A2 [0004]

Claims (10)

Magervormischbrenner an aircraft gas turbine engine with a centrally arranged support burner ( 29 ) and a torch surrounding this annular ( 30 ), characterized in that the supporting burner ( 29 ) a primary fuel channel ( 31 ) and a secondary fuel channel ( 32 ), which can be supplied with fuel independently of each other. A lean burn burner according to claim 1, characterized in that the centric primary fuel channel ( 31 ) is designed for a fuel mass flow rated for igniting the aircraft gas turbine engine. A 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. A lean burn burner according to one of claims 1 to 3, characterized in that the support burner ( 29 ) comprises two concentric swirled pressure atomizers. A lean burn burner according to claim 4, characterized in that the atomizers have a common exit area ( 50 ). A lean burn burner according to one of claims 1 to 3, characterized in that the support burner ( 29 ) one with the primary fuel channel ( 31 ) and a central with the secondary fuel channel ( 32 ) Compressed air atomizer comprises. A lean burn burner according to one of claims 1 to 3, characterized in that the support burner ( 29 ) a centric, with the primary fuel channel ( 31 ) associated pressure atomizer and a surrounding, with the secondary fuel channel ( 32 ) comprises a single jet atomizer. A lean burn burner according to one of claims 1 to 7, characterized in that the primary fuel channel ( 31 ) is arranged centrally and by an annular secondary fuel channel ( 32 ) is surrounded. A lean burn burner according to one of claims 1 to 8, characterized in that the secondary fuel channel ( 32 ) and arranged by an annular primary fuel channel ( 31 ) is surrounded. Method for operating a lean premix burner according to one of Claims 1 to 9, in which a first stage of the support burner (11) for igniting and starting the gas turbine engine ( 29 ) with fuel supply through the primary fuel channel ( 31 ) is applied to a load of the gas turbine of about 20%, in which between a load of about 20% to about 60% of the gas turbine additionally a second stage of the supporting burner ( 29 ) with fuel supply through the secondary fuel channel ( 32 ) and in which between a load of about 60% to 100% of the gas turbine engine, the first stage of the auxiliary burner ( 29 ) and the main burner or the first stage and the second stage of the support burner and the main burner are used.

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