EP1484553B1 - Burner for a gas turbine combustor - Google Patents

Description

The invention relates to a burner for a gas turbine combustor, in particular for an aircraft gas turbine, which comprises a lean burn burner with centrically integrated support burner.

Magervormischbrenner for gas turbine engines and other gas turbines used in the combustion chamber, a fuel-air mixture is burned with high air content and low combustion temperature and correspondingly reduced nitrogen oxide formation, are well known. However, the use of such burners is disadvantageous in that the flame stability is not guaranteed. That is, as the combustion temperature decreases, the air-fuel mixture introduced into the combustion chamber can no longer burn or ignite continuously, so that the flame may fluctuate or go out. In aircraft gas turbine engines, this problem is particularly at low outside temperatures, hail or rain showers or similar unfavorable weather conditions and thereby reduced temperature of the air-fuel mixture. To ignite the air-fuel mixture, a sufficiently high air temperature is required to rapidly vaporize the liquid fuel introduced as droplet mist into the combustion chamber, to preheat, depending on the composition of the fuel-air mixture, the highest possible temperature and thus easier to ignite can.

In order to ensure the ignition of the air-fuel mixture at any time, the arranged in the combustion chamber lean burn burners is known to assign a firing or support burner, with a high combustion temperature is generated in an air-fuel mixture with higher fuel content (rich mixture), to be able to ignite the lean, low combustion temperature producing air-fuel mixture produced by the lean burn burner or main burner even at low air temperature and corresponding unfavorable evaporation behavior of the liquid fuel and to ensure the flame stability.

Usually, the combustion chamber is designed graduated in lean burn burners with support measures, each master / lean burn burner in a lateral gradation in each case a support burner is assigned separately. Apart from the complicated structure, the high number of parts, the high production costs and the high weight, the cooling of the large surfaces is associated with considerable effort. These combustor concepts are commonly referred to as "axially stepped combustors" or "double annular combustors."

The above-described constructive disadvantages occur in another, operated with support measures lean burn burner, in which the pilot burner is centrally integrated into this, not on, but this burner type could not prevail so far, as for a stable operation of the centrally arranged support burner required air-fuel ratio can not be achieved. Particularly critical is the idling operation of the gas turbine, because just then the air inlet temperature into the combustion chamber is low and the startup of the gas turbine after takeoff, because in some cases very high total air-fuel ratios must be traversed. In addition, transient operating points must be mobile: Particularly unfavorable in this sense, the transition from part load under cruise conditions in flight idle for the descent.

In addition, there are maneuver cases in which the engine thrust must be reduced very quickly and set by the decreasing fuel flow extremely lean air-fuel ratios. In addition, as already mentioned, all these unfavorable operating points must also be able to fly under extremely meteorological conditions such as hailstones or tropical rains. Furthermore, conditions must also be able to be realized, such as when the engine is restarted or the combustion chamber is re-ignited at high altitude, i. under atmospheric conditions with very low pressure and low temperature (down to -56 ° C).

A burner combination of the type mentioned above, which has a main burner and a centrally integrated support burner in this, is for example in the EP 0 660 038 B1 described. This burner comprises a main burner with an annular outer fuel-air mixing duct for generating a fuel-air mixture to be introduced into the combustion chamber and a support burner formed in an axially extending channel of a central body, that is centrally accommodated in the main burner, at its discharge opening fuel sprayed and mixed with a core air is introduced into the gas turbine combustor. A stable flame formation in any operating condition can not be achieved with this burner design.

Such a burner is also off EP 1 134 494 known.

The invention has the object of providing a burner of the type mentioned in such a way that over the entire operating range of a gas turbine engine, a stable flame formation in the combustion chamber and an always safe function of the gas turbine is ensured.

According to the invention the object is achieved with a trained according to the features of claim 1 burner for a gas turbine combustor. From the dependent claims, further features and advantageous developments of the invention.

The basic idea of the invention is a lean premix burner with a lean air-fuel mixture supplied via a main air ring channel and a support burner centrally integrated in the lean burn burner with a core air ring channel surrounded by the main air ring channel and a spray nozzle for fuel arranged at the discharge opening therein in that a flame stabilizing ring highly heated by the combustion process is arranged in the adjoining mouth region of the concentrically arranged annular channels, the air guiding flanks of which guide the main air-fuel mixture outwards and the core air flow inwards. The gas flow generated by the hot flame stabilizing ring forms a hot, approximately hollow-cylindrical to barrel-shaped, stationary recirculation zone or hot gas zone which acts together with the stabilizing ring as an ignition element and in which the fuel discharged by the supporting burner is collected and completely burnt , With the flame stabilization ring according to the invention, it is possible to provide a stable, non-extinguishing flame in every operating state of a gas turbine equipped with a lean burn burner and integrated support burner, even when the air temperature is reduced due to external conditions, thus ensuring the functional safety of the gas turbine engine.

According to the invention, the flame stabilizing ring is formed as a conical ring, which has an enclosed by two legs, open towards the combustion chamber groove. On the side facing the burner, the legs form the guide flanks for the inwardly flowing core air or the outwardly flowing air-fuel mixture. At the same time its required cooling against overheating is ensured by the groove or the legs of the conical ring. The cooling takes place at the Luftleitflanken the relatively thin-walled legs of Flame stabilizing ring with the supplied core air or main air.

In an embodiment of the invention, the conical ring (flame stabilization ring) consists of a heat-resistant or heat-resistant or flame-side heat-resistant coated material. The cone ring is mounted with the tip on the face of the central body, which separates the core air ring channel from the main air ring channel.

Fig. 1

eine Schnittansicht eines der Brennkammer einer Flugzeug-Gasturbine zugeordneten Magervormischbrenners mit zentrisch integriertem Stützbrenner; und

Fig. 2

die Brenneranordnung nach Fig. 1, jedoch mit dem Strömungsverlauf der Brennstoffe und der Luft sowie der in der Gasturbinenbrennkammer ausgebildeten Heißgas- oder Rezirkulationszone.

An embodiment of the invention will be explained in more detail with reference to the drawing. Show it:

Fig. 1

a sectional view of the combustion chamber of an aircraft gas turbine associated lean premix burner with centrally integrated support burner; and

Fig. 2

the burner assembly according to Fig. 1 but with the flow of the fuel and air and the hot gas or recirculation zone formed in the gas turbine combustor.

The burner 1 has a housing 2 and a central body 3 with a main air annular channel 4 formed between them for a main or lean burn burner, which is assigned to a combustion chamber 5 of a (aircraft) gas turbine. In the main air-annular channel 4 of the lean burn burner through which flow about 90% of the total burner air, main air swirl generator 6 are involved, which put the main air flow - arrow A - in a rotary motion. In the main air flow liquid fuel is injected after the swirl generation, which mixes with the hot air stream and partially evaporated in this. The introduced into the combustion chamber 5 - lean - fuel-air mixture has a high proportion of air and burns in the combustion chamber 5 with a correspondingly low combustion temperature, so that the nitrogen oxide emission and consequently the air pollution is extremely low.

The low pollutant emission achieved with low combustion temperatures, on the other hand, however, is associated with a flame instability occurring in particular in unfavorable weather conditions due to the consequent reduced air inlet temperature.

To ensure a safe formation of the flame, for example, for rapid acceleration and deceleration of the gas turbine, and to avoid extinguishing the flame in the central body 3 is a running in the central axis channel 7 for receiving a support burner is formed, comprising a nebulizer, consisting of atomizer fins 18, a fuel line 8, a subsequent thereto atomizer carrier tube 9 and an opening into the combustion chamber 5 atomizer nozzle 10, and formed on the circumference of the atomizer core air ring channel 11 comprises. Core air supplied in accordance with the arrow B passes through the core air annular channel 11 and a core air swirler 12 for generating axial rotational movement of the core air into the gas turbine combustor 5 to produce a fuel air therein to produce a stable flame with the fuel sprayed from the atomizer nozzle 10 To provide a mixture with a high fuel content. The direction of rotation of the main air flow and the core air flow is preferably in the same direction.

An essential component of the present lean burn burner with centrally integrated auxiliary burner is a flame stabilization ring adjoining the central body 3 in the mouth region of the core air ring channel 11 and the main air ring channel 4, which is in the form of a conical ring 13 with a tip adjoining the central body 3 and one of an annular core air flank 14 and an annular Hauptluftleitflanke 15 formed, facing the interior of the combustion chamber 5 groove 16 is executed. By the Kernluftleitflanke 14 inwardly deflected core air flow and generated with the Hauptluftleitflanke 15, outwardly directed main air flow is formed in the combustion chamber 5 outgoing from the groove 16, formed substantially hollow cylindrical and curved, stationary recirculation zone 17 with the highest temperature (hot gas zone), that is, a stable flame area with in the flute 16 lying flame root, which compensate in the recirculation zone 17, the velocities of the currents generated by the main air ring channel 4 and the core air ring channel 11. In this dormant, hot recirculation zone 17, the cold air supply due to unfavorable weather conditions so far not evaporated fuel mist that penetrate from the atomizer 10 in this area, or fuel-air mixtures at high temperature dwell sufficiently long and evaporate as much as possible to a good burn - To form and ignitable fuel-air mixture in the combustion chamber. The fuel exit angle from the atomizer nozzle 10 is adjusted so that the fuel droplets meet the very hot, stationary recirculation zone 17 and are burned there and can not reach the combustion chamber walls beyond this area.

To the formation of the curved hollow cylindrical hot recirculation zone 17 contributes significantly to the heating of the groove 16 having the conical ring 13, at its due to the lying there flame root hot surface of the fuel or the fuel-air mixture is also ignited to maintain the combustion. The conical ring 13 is made of heat-resistant steel, optionally with a ceramic protective coating on the flame side, or even completely of a ceramic material (advantageous: fiber-ceramic composite materials). Overheating of the conical ring 13 is prevented by suitable choice of material and the good heat transfer to the relatively thin-walled core and Hauptluftleitflanken 14, 15 of the conical ring 13 and at the rear as a cooling medium flowing main air (air-fuel mixture) or core air.

LIST OF REFERENCE NUMBERS

1

burner

2

casing

3

central body

4

Main air ring channel

5

Gas turbine combustor

6

Home Air swirler

7

channel

8th

fuel line

9

Nebulizer support tube

10

atomizer

11

Core air ring channel

12

Core air swirler

13

Cone ring (flame stabilizer ring

14

Kernluftleitflanke

15

Hauptluftleitflanke

16

fillet

17

Recirculation zone, hot gas zone

18

Zerstäuberflossen

Arrow A.

Main air flow, air-fuel mixture

Arrow B

Core airflow

Claims (4)

1. Burner (1) for the combustion chamber (5) of a gas turbine, in particular an aircraft gas turbine, with centrally integrated stabilizing burner,
   where a core air annulus (11) accommodating an atomizer nozzle (10) of the stabilizing burner (20) is surrounded by a main air annulus (4), with the atomizer nozzle (10) issuing into the combustion chamber (5),
   where the fuel discharge angle of the atomizer nozzle (10) is between 60° and 130° and preferably amounts to 95°,
   where main air or core air swirlers (6 and 12, respectively) are arranged in the main air and the core air annulus (4, 11),
   where in the adjacent issuing areas of the main air and the core air annulus (4, 11) a flame stabilizing ring (13) designed as conical ring is arranged, whose apex connects to a central body (3) which separates the core air annulus (11) from the main air annulus (4), and characterized in that the conical ring is provided with an inwardly directed core air deflector flank (14) for inwardly deflecting the core air flow (arrow B) and with an outwardly directed main air deflector flank (15) for outwardly deflecting the main air flow (arrow A) and has a fillet (16) on the side facing the gas-turbine combustion chamber (5),
   where a lean air fuel mixture is supplied to the burner (1) via the main air annulus (4) and a fuel air mixture with high fuel content via the core air annulus (11) and the atomizer nozzle (10) of the stabilizing burner (20) and where, originating at the fillet (16), an essentially hollow-cylindrical and cambered, stationary, hot recirculation zone (17) is formed as a stable flame area, extending with maximum temperature into the gas-turbine combustion chamber (5), with the flame root being situated in the fillet (16).
2. Burner in accordance with Claim 1, characterized in that the conical ring (21) is made of heat-resisting steel.
3. Burner in accordance with Claim 1 or 2, characterized in that the conical ring (21) is provided with a ceramic coating in the area of the fillet (16).
4. Burner in accordance with Claim 1, characterized in that the conical ring (21) is made of ceramic material.

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