US20170299183A1

US20170299183A1 - Method for operating a lean premix burner of an aircraft gas turbine and device for carrying out the method

Info

Publication number: US20170299183A1
Application number: US13/974,687
Authority: US
Inventors: Imon Kalyan BAGCHI; Waldemar Lazik
Original assignee: Rolls Royce Deutschland Ltd and Co KG
Current assignee: Rolls Royce Deutschland Ltd and Co KG
Priority date: 2012-08-28
Filing date: 2013-08-23
Publication date: 2017-10-19
Also published as: EP2703720A2; DE102012017065A1; EP2703720A3

Abstract

The present invention relates to a method for operating a lean premix burner of an aircraft gas turbine, where fuel and primary supporting air are supplied by means of a supporting burner (pilot burner) arranged centrically to the burner axis, where secondary air surrounding the supporting burner is supplied, and where fuel and air are supplied by means of a main burner, characterized in that the primary supporting air is supplied in an amount of 5 vol % to 10 vol % of the total air quantity, that the secondary supporting air is supplied in an amount of 5 vol % to 20 vol % and that 35 vol % to 75 vol % of the total air quantity are supplied via the main burner in the partial load range and in the full load range.

Description

This invention relates to a method for operating a lean premix burner of an aircraft gas turbine, where fuel and primary supporting air are supplied by means of a supporting burner (pilot burner) arranged centrically to the burner axis and where fuel and air are supplied by means of a main burner.
It is known from the state of the art to use two fuel atomizers, i.e. a supporting burner and a main burner, in lean premix burners. The supporting burner is arranged centrically in the main burner. The supporting burner is here usually designed as a pressure swirl atomizer. The lean premix burner includes here two fuel lines for supplying the supporting burner and the main burner. In operation, the supporting burner is used for igniting the aircraft gas-turbine engine and in low load conditions, whereas the main burner is put into operation at partial load and is used up to maximum load. The supporting burner is designed here for the ignition operation and for a stable combustion during the engine starting phase.
The state of the art is described in the following in light of FIG. 2, which shows a burner 32 arranged on a combustion chamber head 31 and supplying the combustion chamber with fuel and approximately 10 vol % to 20 vol % of the total air (strictly speaking this is mass %, but in this case they are identical since the air has a constant temperature.) As a result, a rich zone 33 is formed, which is arranged directly downstream of the burner 32. A further 30 vol % to 40 vol % of air are supplied through mixing air openings 34 to 37. This results in an air admixture to the rich flame in the flame zone 38. Downstream of this flame zone 38, a lean zone 39 is provided. The remaining air of 40 vol % to 50 vol % is used for cooling and flows through an inner combustion chamber wall 40 and an outer combustion chamber wall 41, which maintain the flame. FIG. 2 thus shows a standard burner with a rich zone supplied with air and followed by a lean zone.
FIG. 3 shows an embodiment according to the state of the art, where the burner 32 includes means for mixing air and fuel. A direct or further flame zone 38 as shown in FIG. 2 can be dispensed with. The entire burner (supporting burner and main burner) passes 50 vol % to 80 vol % of the total air into the combustion chamber. The remaining air quantity of 20 vol % to 50 vol % is used for cooling. The burner includes two fuel circuits and thus permits the supply of fuel through two concentric fuel atomizers. The supporting burner 42 with the associated atomizer supplies 5 vol % to 15 vol % of the total air and creates a small rich zone 33 which is used for starting the engine and for flame stability. The concentric main fuel atomizer 43 supplies fuel at medium to maximum load conditions and 40 vol % to 75 vol % of the total air 44. This creates a lean zone 39 surrounding the rich zone 33. This lean zone 39 is responsible for low pollutant emissions, in particular of NOx.
The main drawback of the solution shown in FIG. 2 is that high pollutant emissions result, in particular of NOx, and that in high load conditions soot is emitted, since the combustion conditions approximate to a stoichiometric or rich combustion state. It was therefore attempted in the solution described in FIG. 3 to optimize the combustion process by a lean burner concept. This however has the disadvantage that the supporting burner (pilot burner) has a reduced flame stability at low output of the aircraft gas turbine. In medium load conditions, the combustion by the main burner is too lean to operate effectively, leading to increased fuel consumption by an aircraft. Additionally, the absence of high air admixing results in low oxidation of soot, so that considerable soot quantities are emitted from the aircraft gas turbine at medium load.
The object underlying the present invention is to provide a method—and a device for carrying out the method—for operating a lean premix burner which avoid the disadvantages of the state of the art and enable, in particular, a good, stable and low-pollutant combustion.
It is a particular object of the present invention to provide solution to the above problematics by a combination of the features of the independent Claims. Further advantageous embodiments of the present invention become apparent from the sub-claims.
It is thus provided in accordance with the invention that between the rich zone and the lean zone an intermediate admixing zone with a high jet-like admixture of air is formed, into which zone an additional fuel/air flow is introduced. This results in optimized combustion in partial load areas too, and has the advantage that the soot emissions are reduced by improved oxidation of the soot. Furthermore, there is improved combustion with better efficiency, since the very lean zones known from the state of the art are avoided. Due to the intermediate admixing zone, a combustion zone is created which is closer to stoichiometric fuel/air ratios. Although this zone is still lean, it avoids the disadvantages of a too-lean combustion zone.
In accordance with the invention, there is an enrichment of the rich zone with a lower air proportion by the supporting burner. Instead, the additional air is introduced into the intermediate admixing zone. This leads to good flame stability and good ignitability of the aircraft gas turbine.

The present invention is described in the following in light of the accompanying drawing, showing exemplary embodiments. In the drawing,

FIG. 1 shows a schematic representation of a gas-turbine engine in accordance with the present invention,

FIG. 2 shows a longitudinal sectional view of a combustion chamber in accordance with the state of the art,

FIG. 3 shows a schematic representation of a further variant of a combustion chamber in accordance with the state of the art by analogy with FIG. 2,

FIG. 4 shows a simplified longitudinal sectional view of a combustion chamber in accordance with a first exemplary embodiment of the invention by analogy with the representation of FIG. 3,

FIG. 5 shows a representation of a further exemplary embodiment by analogy with FIG. 4,

FIG. 6 shows a sectional view of a further exemplary embodiment by analogy with FIGS. 4 and 5,

FIG. 7 shows an enlarged partial representation of the flow conditions of the inventive solution in accordance with FIG. 6,

FIGS. 8 to 11 show sectional, front and perspective views of differing exemplary embodiments of flame stabilizers and secondary air recesses,

FIG. 12 shows a graphic representation of the equivalence ratio as a function of the thrust in accordance with the state of the art,

FIG. 13 shows a graphic representation of a lean premix burner, by analogy with FIG. 12, and

FIG. 14 shows a representation of the inventive solution by analogy with FIGS. 12 and 13.

The gas-turbine engine 10 in accordance with FIG. 1 is a generally represented example of a turbomachine where the invention can be used. The engine 10 is of conventional design and includes in the flow direction, one behind the other, an air inlet 11, a fan 12 rotating inside a casing, an intermediate-pressure compressor 13, a high-pressure compressor 14, a combustion chamber 15, a high-pressure turbine 16, an intermediate-pressure turbine 17 and a low-pressure turbine 18 as well as an exhaust nozzle 19, all of which being arranged about a center engine axis 1.
The intermediate-pressure compressor 13 and the high-pressure compressor 14 each include several stages, of which each has an arrangement extending in the circumferential direction of fixed and stationary guide vanes 20, generally referred to as stator vanes and projecting radially inwards from the engine casing 21 in an annular flow duct through the compressors 13, 14. The compressors furthermore have an arrangement of compressor rotor blades 22 which project radially outwards from a rotatable drum or disk 26 linked to hubs 27 of the high-pressure turbine 16 or the intermediate-pressure turbine 17, respectively.
The turbine sections 16, 17, 18 have similar stages, including an arrangement of fixed stator vanes 23 projecting radially inwards from the casing 21 into the annular flow duct through the turbines 16, 17, 18, and a subsequent arrangement of turbine blades 24 projecting outwards from a rotatable hub 27. The compressor drum or compressor disk 26 and the blades 22 arranged thereon, as well as the turbine rotor hub 27 and the turbine rotor blades 24 arranged thereon rotate about the engine axis 1 during operation.
FIG. 4 shows in a schematic representation a longitudinal sectional view of a burner in accordance with the invention. It includes a supporting burner 42 and a fuel atomizer 43 surrounding the latter, both forming part of a burner 32 mounted on a combustion chamber head 31. The combustion chamber includes an inner combustion chamber wall 40 and an outer combustion chamber wall 41. Air and fuel are supplied by the supporting burner 42 for forming a rich zone 33 immediately adjoining the supporting burner 42. In total, an air quantity of approx. 50 vol % to 80 vol % of the total air is supplied to the combustion chamber by the burner. The fuel is supplied via two concentric atomizers. Only a small amount of air (5 vol % to 10 vol % of the total combustion chamber air) is supplied via the atomizer of the supporting burner 42, thus ensuring that the rich zone 33 is created.
The rich zone 33 is delimited and partially enclosed by an intermediate admixing zone 45. An air quantity of 5 vol % to 20 vol % of the total combustion air of the combustion chamber is introduced into the intermediate zone in order to provide a second zone or secondary supporting zone (intermediate admixing zone) 45 forming a further admixing zone (quenching zone).
The main fuel atomizer 43 supplies fuel and air. The air quantity supplied is 35 vol % to 75 vol % of the total combustion chamber air. The main fuel atomizer 43 is used during medium to maximum load states of the aircraft gas turbine. By supplying air and fuel via the main fuel atomizer 43, a lean zone 39 is created which surrounds the intermediate admixing zone 45 and adjoins the latter in the axial direction (flow direction).
FIG. 6 shows a detailed representation of a further exemplary embodiment of the invention, by analogy with the representation in FIG. 4. Identical parts are provided with the same reference numerals, as is the case in the following exemplary embodiments.
FIG. 5 shows in detail the supporting burner 42 with a fuel outlet 47. The supporting burner is concentrically surrounded by an annular air passage 48, in which a swirler element 49 is arranged. The escaping air/fuel mixture creates the rich zone 33.
The intermediate admixing zone 45 is formed by the further supply of air and fuel. A concentric annulus 50 is provided for this. The design permits a greater pressure drop, in order to generate higher air velocities at the place where air is introduced into the combustion chamber. This results in good mixing with the rich zone 33. The secondary air supply 51, 52 and 53 can take place through suitable recesses described in the following in conjunction with FIGS. 8 to 10.
The main fuel is supplied through a concentric main air supply 54 and atomized by the inner air supply 55 and mixed with the latter. A swirl is imparted by an inner main swirler element 56. The main fuel is also guided through an outer air supply 57 and atomized and mixed with it, with a swirl being imparted to this air supply by means of an outer main swirler element 58. The flame resulting from the main burner surrounds the intermediate admixing zone 45 and forms a lean zone 39.
The secondary air can be supplied at different points ( secondary air supply 51, 52 or 53). This supply can take place singly or in combination.
FIG. 11 shows an axial longitudinal sectional view plus a front view of an exemplary embodiment of a burner in accordance with the invention. It is shown here that the secondary air recesses 52 can be designed in the form of round holes provided on a flame stabilizer.
In the exemplary embodiment in FIG. 9, the secondary air is supplied by tubes (chutes) provided on the flame stabilizer 59. It can be supplied in either an axial or a tangential alignment in order to impart a swirl to the secondary air. Between 4 and 36 of these outlet tubes (chutes) can be provided, being at angles of 0° and 60° to the burner axis.
FIG. 11 shows a further design variant in which the secondary air recesses 52 are designed in the form of slots. Between 4 and 36 slots can be provided, and can have an angle between 0° and 60° relative to the burner axis in order to impart an additional swirl to the air.
FIG. 8 shows a further exemplary embodiment with V-shaped slots 52, which can also be provided in a number between 4 and 36. Here too it is possible to incline the V-shaped slots relative to the burner axis between 0° and 60° for further swirling of the air.
The burner described above can also be designed with an onflow supporting burner, as is shown in FIG. 6. The supporting fuel is supplied through a fuel outlet 47. The supporting air is supplied through an inner air passage 48 with a swirler element 49 and an outer annulus 50 with a swirler element 56.
FIG. 7 shows that in accordance with the invention a second supporting flame stabilization zone Y is formed additionally to zone X and to the rich zone 33. The zone Y leads to an improved interaction between the supporting burner and the main burner. In certain operating conditions, the main flame can also be stabilized in zone Y.
In accordance with the invention, an additional intermediate zone is thus created by which combustion in the combustion chamber can take place in a controlled and optimized way. This leads to the supporting burner zone being able to operate in a stable manner, without any fear of the supporting burner being extinguished. The intermediate zone can be operated even in relatively high load conditions without soot emissions. Furthermore, the intermediate admixing zone improves combustion efficiency (total combustion) during staged operation of the main burner. This leads to a minimum drop in the efficiency of combustion during a transition from operation of the supporting burner to combined operation of the supporting burner and of the main burner.
In the following, the invention is again explained in respect of the method in accordance with the invention in light of FIG. 14, where the illustrations in FIGS. 12 and 13 reflect the underlying state of the art.
FIG. 12 shows a diagram in which the thrust is plotted in percentages as a function of the equivalence ratio between air and fuel. With an equivalence ratio of 1, there is a stoichiometric ratio, below 1 to 0 the result is a rich combustion, while above 1 a lean combustion is obtained. These illustrations are also shown in FIGS. 13 and 14.
Furthermore, the information for the combustion zones relates to FIGS. 2 to 5.
FIG. 12 shows an illustration from the state of the art which has high emission values. In particular, at high thrust or high output, respectively, the NOx values are high and there is a lot of soot. The respective equivalence ratios of the individual zones achieve here, as is indicated in FIG. 12 for the zones 33, 38 and 39, values having an equivalence ratio of 1 or a rich equivalence ratio close to the stoichiometric value. By contrast, the result for the supporting burner is good flame stability.
To avoid the drawbacks of increased soot formation and high NOx emissions, solutions were proposed as shown in FIG. 13. While FIG. 12 in particular relates to the representation in FIG. 2, the values shown in FIG. 13 are based especially on an embodiment in accordance with FIG. 3. As shown in FIG. 13, the supporting burner is operated somewhat more leanly. This leads to a good combustion, but at the same time generates a lot of soot. At the same time a reduced stability at low load results from the leaner supporting burner. As also shown in FIG. 13, the burner is set very lean at medium thrust, so that in this partial load area or transition area there is no good combustion in particular in zone 39. A low efficiency thus applies, and this leads to increased fuel consumption of an aircraft.
Furthermore, the absence of the flame zone 38 and the poor interaction between the mixing air 36 lead to a poor oxidation of soot, resulting in high soot emissions.
The drawbacks of the mode of operation shown in FIG. 13 can be reduced according to the state of the art in that the supporting burner is enlarged to pass a larger air quantity through the supporting burner. Soot formation could be reduced by this, but this has the negative effect that higher NOx emissions result. Moreover, a leaner operation of the supporting burner leads to a lower stability. Furthermore, a second supporting burner circuit with a total of three fuel circuits could be introduced, but this would increase the complexity of the overall system, involving additional costs for fuel injection nozzles, fuel systems and control systems.
Based on the procedures described above, a completely different solution was created in accordance with the invention, and is explained in light of FIG. 14.
The solution in accordance with the invention was described above in particular in conjunction with the design solution according to FIG. 4.
In accordance with the invention, a secondary supporting zone or intermediate admixing zone 45 is formed, as explained above, which is achieved by diverting air/fuel from the rich zone 33. Furthermore, there is a diversion of fuel and air from the total airflow 44. By doing so, an additional flow 46 is used, as is shown in FIG. 4.
The solution in accordance with the invention results in the following advantages:
As shown in FIGS. 4 and 5, the addition of the secondary supporting zone/intermediate admixing zone 45 leads to a reduction in the soot emissions, caused by an improved oxidation of the soot. Furthermore, there is an improved combustion efficiency due to the reduction of very lean zones. The zone of the main burner remains lean, but zone 45 forms a secondary supporting zone or intermediate admixing zone which is closer to the stoichiometric fuel/air ratio. Furthermore, the zone 45 leads to a reduction in soot formation. The zone 33 (zone of the supporting burner 42) can be operated with a richer fuel/air mixture than in the solutions known from the state of the art. This leads to an improved flame stability.
As explained above, the core of the invention is the additional introduction of a secondary supporting zone or intermediate admixing zone 45. This leads to the supporting burner being capable of operation with a constant combustion zone, thereby ensuring stable operation and preventing the flame from being extinguished (flame-out). Both the supporting burner zone and the secondary supporting zone/intermediate admixing zone 45 can be operated without problems resulting with regard to soot emissions. Furthermore, the secondary supporting zone/intermediate admixing zone 45 improves the efficiency of combustion during a staged operation of the main burner. This leads to a minimum reduction in the combustion efficiency during the transition from operation with the supporting burner to combined operation of the supporting burner and of the main burner.

LIST OF REFERENCE NUMERALS

1 Engine axis
10 Gas-turbine engine/core engine
11 Air inlet
12 Fan
13 Intermediate-pressure compressor (compressor)
14 High-pressure compressor
15 Combustion chamber
16 High-pressure turbine
17 Intermediate-pressure turbine
18 Low-pressure turbine
19 Exhaust nozzle
20 Guide vanes
21 Engine casing
22 Compressor rotor blades
23 Stator vanes
24 Turbine blades
26 Compressor drum or disk
27 Turbine rotor hub
28 Exhaust cone
31 Combustion chamber head
32 Burner
33 Rich zone
34, 35, 36, 37 Mixing air
38 Flame zone
39 Lean zone
40 Inner combustion chamber wall
41 Outer combustion chamber wall
42 Supporting burner
43 Fuel atomizer
44 Total air
45 Secondary supporting zone/intermediate admixing zone
46 Additional flow
47 Fuel outlet
48 Air passage
49 Swirler element
50 Annulus
51, 52, 53 Secondary air supply/secondary air recesses
54 Concentric main air supply
55 Inner air supply of main burner
56 Inner main swirler element
57 Outer air supply of main burner
58 Outer main swirler element
59 Flame stabilizer
60 Supporting air supply

Claims

1. Method for operating a lean premix burner of an aircraft gas turbine, where fuel and primary supporting air are supplied by means of a supporting burner (pilot burner) arranged centrically to the burner axis, where secondary air surrounding the supporting burner is supplied, and where fuel and air are supplied by means of a main burner, characterized in that the primary supporting air is supplied in an amount of 5 vol % to 10 vol % of the total air quantity, that the secondary supporting air is supplied in an amount of 5 vol % to 20 vol % and that 35 vol % to 75 vol % of the total air quantity are supplied via the main burner in the partial load range and in the full load range.

2. Method in accordance with claim 1, characterized in that adjacent to the supporting burner a rich zone is formed, that the rich zone is enclosed by an intermediate admixing zone, that the intermediate admixing zone is enclosed and that the intermediate admixing zone is enclosed by a lean zone.

3. Aircraft gas turbine lean premix burner for carrying out the method in accordance with claim 1, characterized in that a flame stabilizer concentrically surrounding the supporting burner is provided with secondary air recesses.

4. Premix burner in accordance with claim 3, characterized in that the secondary air supply recesses are provided in the form of straight or V-shaped slots.

5. Premix burner in accordance with claim 3, characterized in that the secondary air supply recesses are provided in the form of tubes (chutes).