JPH102558A - Fuel nozzle for gas turbine combustor - Google Patents

Abstract

(57) [Summary] (Problem corrected) [Problem] To cool the burner end face without reducing the ratio of combustion air to the total air introduced into the combustor from the air compressor in the full load band, and achieve low NOx. Provide a highly reliable gas turbine combustor. A passage, which is located at an upstream end of a combustion chamber of a gas turbine and guides combustion air to the combustion chamber, and a first swirler vane, located in the passage, which swirls the combustion air.
And a fuel nozzle for ejecting a fuel gas into the passage downstream of the first swirling blade or the first swirling blade. As a result, the cooling space provided in the burner end face portion 18 is guided to the cooling space, and further discharged through the second swirling blade 25 provided in the outer periphery of the cooling space to the passage downstream of the first swirling blade.

Description

DETAILED DESCRIPTION OF THE INVENTION

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a gas turbine combustor in which the cooling performance of a burner end face member in a gas turbine combustor is enhanced and the reliability of the burner is improved.

[0002]

2. Description of the Related Art An example of the overall structure of a gas turbine combustor will be described with reference to FIG.

First, high-pressure air 1 compressed by an air compressor (not shown) flows in an annular flow path between an outer cylinder 2 and a combustion cylinder 3 of a combustor in the opposite direction, as shown by arrows 4 and 5. Into two streams. As indicated by arrow 6, the flow of arrow 4 flows in the annular passage 7 after flowing in the direction of the central axis of the combustor, flows out into the combustion chamber 9 while receiving the swirling force by the swirling blade 8, and is ejected from the fuel injection hole 10. A pilot flame 11 is formed in the combustion chamber 9 together with the fuel. On the other hand, the high-pressure air indicated by the arrow 5 is mixed with the fuel ejected from the premixed fuel nozzle 12, and after the degree of mixing of the fuel and the air is increased in the premixer 13, the flame stabilizer 14 enters the combustion chamber. A premixed flame 15 is formed.

FIG. 2 shows the state of the flow near the pilot flame. The air-fuel mixture introduced into the combustion chamber 9 from the air-fuel mixture outlet 16 has a swirl component due to the swirl vanes 8.
Due to the centrifugal force, the flow spreads radially outward, so that the pressure on the central axis of the combustion chamber is reduced, and a pressure gradient is generated on the central axis in a direction opposite to the flow of the main flow, and a recirculating flow 17 is generated. The gas temperature in the region of this recirculation flow is very high and the burner end face 18 in contact with this region
Since the temperature of the burner is also high, some measures must be taken to cool the burner end face in order to improve the reliability of the combustor.

A method of cooling a burner nozzle for a gas turbine combustor is exemplified in, for example, Japanese Patent Application Laid-Open No. 5-172331.

In the first example, a forced convection cooling by air flowing in the flow path is provided by providing a flow path that communicates a portion of the combustion air from the upstream side of the swirling vanes to the end face of the fuel injection nozzle and directly flowing out to the combustion chamber. The structure has an effect and a film cooling effect by air flowing out to the nozzle end face.

[0007] The second example is an example in which fuel gas is used as a cooling medium. A structure is used in which forced convection cooling is performed by the impinging jet by creating a flow in which the fuel gas collides with the end face of the nozzle in the burner nozzle. Has become.

[0008]

In recent years, with the trend toward higher output and higher efficiency in gas turbine plants, the combustion gas temperature of the gas turbine combustor has been increasing.

[0009] Therefore, in order to suppress the generation amount of nitrogen oxide (NOx) as an air pollutant, which is known to increase with the rise of the combustion gas temperature, various combustion systems have been proposed. Although adopted in turbine combustors, from a comprehensive point of view, the fuel and air are mixed beforehand, and the lean reaction of increasing the ratio of the air flow rate to the fuel flow rate and keeping the combustion gas temperature low while proceeding with the combustion reaction. Many premixed combustion methods are used.

[0010] However, for this purpose, as much as possible, the limited air introduced into the combustor from the air compressor is used as combustion air to dilute the premixed air and to expose the combustion wall exposed to high temperatures. I want to keep the ratio of cooling air for cooling such as low as possible. Therefore, the method of using a part of the air only for cooling the nozzle end face as shown in the first example of the prior art has a problem in view of the lean premixed combustion system.

Further, in such a structure, by extracting a part of the combustion air, the flow rate of the air introduced into the swirler is reduced by that amount, and the flow rate of the air-fuel mixture at the outlet of the swirler is reduced. Therefore, when the ratio of the fuel flow rate to the air flow rate (hereinafter referred to as the fuel-air ratio) is increased, the speed at which combustion proceeds becomes faster than the flow rate of the air-fuel mixture at the swirl vane outlet. There is a risk that a flame may enter the inside and burn the burner.

Regarding the cooling method shown in the second example,
For example, in a gas turbine combustor in which this burner is arranged as a pilot burner at the center of the upstream end of the combustion chamber and a premix burner for performing premix combustion is arranged on the outer peripheral side thereof, when the gas turbine output is low. If the pilot burner is used for combustion alone and the gas turbine output is increased, the following precautions will be taken if the premix burner on the outer periphery is ignited to increase the fuel-to-air ratio. .

Generally, in diffusion combustion, the amount of nitrogen oxides generated in the combustion gas increases because the local combustion gas temperature increases, and as described above, the premix combustion reduces the combustion gas temperature. It is known that the amount of nitrogen oxides generated is small in order to promote the combustion reaction while keeping it low, and from this fact it is also necessary to keep the amount of nitrogen oxides low under rated conditions with a high combustion load. It is necessary to reduce the diffusion combustion in the pilot burner as much as possible and adopt a combustion method mainly using the combustion in the premix burner.

[0014] However, in the cooling method shown in the second example, the fuel flow rate to the pilot burner decreases despite the increase in the gas turbine output and the gas temperature in the combustion chamber. The fuel flow rate is reduced, the cooling effect is reduced, and the nozzle may be burned out.

An object of the present invention is to cool the burner end face without reducing the ratio of combustion air to the total air introduced into the combustor from the air compressor in the full load zone, and achieve a low NO.
x to provide a highly reliable gas turbine combustor.

[0016]

The first feature of the present invention is as follows.
Partially bleeding the air or air-fuel mixture in the passage that guides the air or air-fuel mixture to the combustion chamber, guides it to the cooling space provided at the nozzle end face, and performs forced convection cooling of the nozzle end face,
While flowing through the second swirling blade provided on the outer periphery of the nozzle end face, the swirling force flows out of the first swirling blade downstream in the passage.

A second feature of the present invention is that a passage portion for introducing air or an air-fuel mixture into the combustion chamber partially extracts the air or the air-fuel mixture and guides the air or the air-fuel mixture to an end face of the nozzle. In this case, the nozzle end face is forcedly cooled by convection cooling by flowing through the flow path that extends, and flows out of the passage downstream of the first swirl vane.

According to the present invention, the air or the air-fuel mixture is partially extracted from the main flow in the passage for guiding the air or the air-fuel mixture to the combustion chamber, and is led to the cooling space provided at the nozzle end face. The nozzle end face is subjected to forced convection cooling, whereby the reliability of the nozzle member can be improved.

The air or air-fuel mixture cooled at the nozzle end face is not directly discharged to the combustion chamber, but directly or through a second swirler provided on the outer periphery of the nozzle end face, while having a swirl component and downstream of the first swirler. And merges again with the main flow, so that the flow speed of the main flow into the combustion chamber is suppressed from being reduced, and as a result, the flow speed is lower than the speed at which combustion proceeds, and the flame is reduced by the first swirl blade. While preventing the first swirl vane from burning out
The cooling of the nozzle member becomes possible.

[0020]

FIG. 3 to FIG. 5 show a first embodiment of the present invention.

FIG. 5 is a view of the burner of this embodiment as viewed from the downstream direction. FIGS. 3 and 4 are cross-sectional views taken along line AA of FIG.
It shows a BB cross section.

The burner includes an air flow path 19, a cooling air inlet hole 20, a cooling air passage 21, a fuel gas inlet hole 22, a fuel pipe 23, a fuel ejection hole 24, a swirling blade 8, and a swirling blade 25.
It is composed of

First, air 6 introduced from an air compressor (not shown) is divided into a flow 26 entering the first swirl vane 8 and a flow 27 entering the cooling air inlet hole 20 while passing through the air flow path 19. The gas 28 is injected from the fuel gas inlet hole 22 through the fuel pipe 23 into the air flow 29 from the fuel outlet hole 24 downstream of the swirl vanes 8 and mixed, and is introduced into the combustion chamber 9 from the mixture outlet 30 to form a flame. It becomes high temperature combustion gas.

Here, the cooling air 27 extracted from the combustion air 6 flows through the cooling air flow path 21, and then is subjected to forced convection cooling 32 at the nozzle end face in the cooling space 31.
Burnout of the burner end face portion 18 can be prevented, and the reliability of the burner can be improved.

The cooling air 32 receives the turning force from the turning blade 8 and the turning blade 25 of the forward turning or the reverse turning,
The cooling air 32 is again introduced into the combustion chamber 9 after being merged with the mainstream air-fuel mixture. It is possible to suppress a decrease in the outflow speed of the air-fuel mixture at the air outlet portion 30, and the fuel flow rate can be reduced by two or more compared to the conventional cooling method.
Even when the load is increased by increasing the load 8, it is possible to prevent the outflow speed of the air-fuel mixture from becoming lower than the speed at which the combustion proceeds, and to prevent the flame from entering the swirl blade 8 and burning the swirl blade.

Further, the cooling air passage 2 is formed in the burner end face 18.
It is also possible to provide a plurality of radially extending flow passages communicating with 1 and to flow cooling air to cool the burner end face portion 18.

Next, a second embodiment of the present invention is shown in FIGS.

FIG. 8 is a view of the burner of this embodiment as viewed from the downstream direction, and FIGS. 6 and 7 are sectional views taken along lines CC and DD of FIG. 8, respectively.

The burner includes an air passage 19, a premixed fuel nozzle 33, a cooling mixture inlet 34, and a cooling mixture passage 3.
5, a diffusion fuel inlet hole 36, a diffusion fuel pipe 37, a diffusion fuel ejection hole 38, a swirl blade 8, and a swirl blade 25.

First, air 6 introduced from an air compressor (not shown) is mixed with premixed fuel 39 ejected from a premixed fuel nozzle 33 provided in the air flow path 19, and fuel
The premixed air is divided into a flow 40 entering the first swirl vanes 8 and a flow 41 entering the cooling mixture inlet 34.

The latter flow 41 is guided to the burner end face 18 through the cooling mixture passage 35 and performs forced convection cooling 43 in the cooling air 42 provided at the burner end face.
The air-fuel mixture 44 flows out of the swirling blade 8 while being swirled by the swirling blade 8 and the forward or backward swirling blade 25 provided on the outer peripheral portion of the burner end face, and is guided to the combustion chamber 9. Further, the diffusion fuel 45 is provided in the diffusion fuel inlet hole 3.
6, the fuel enters the diffusion fuel pipe 37, is drawn out from the diffusion fuel ejection hole 38 into the combustion chamber 9, forms a stable pilot flame (diffusion flame), and flows out from the swirling vanes 8 into the combustion chamber 9 using the pilot flame as a fire. The air-fuel mixture performs premix combustion.

According to the second embodiment, even in a burner in which the combustion performs stable diffusion combustion at the center of the burner and performs the premix combustion using a swirler or a flame stabilizer at the outer periphery thereof, The burner end face can be cooled without lowering the flow velocity at the air outlet, and the burner combustion load can be increased more than before, while improving the reliability of the burner.

In the second embodiment, a plurality of radially extending flow paths communicating with the cooling air passage 35 may be provided in the burner end face portion 18 to flow cooling air to cool the burner end face portion.

[0034]

According to the present invention, the burner nozzle end face which is exposed to the combustion gas and has a high temperature without decreasing the ratio of the combustion air to the total air flow introduced into the combustor from the air compressor. The cooling of the portion becomes possible, and the reliability of the nozzle member can be improved. In other words, it is possible to increase the combustion load of the combustor while using lean combustion that emits a small amount of nitrogen oxide (NOx), which is an air pollutant.

[Brief description of the drawings]

FIG. 1 is an explanatory diagram of an example of the overall configuration of a gas turbine combustor.

FIG. 2 is an explanatory diagram of a flow near a pilot burner in FIG. 1;

3 is a sectional view of the burner of the present invention shown in FIG.

4 is a sectional view of the burner of the present invention shown in FIG. 5, taken along line BB.

FIG. 5 is an explanatory view of the burner of the first embodiment of the present invention viewed from a downstream direction.

6 is a sectional view of the burner of the present invention shown in FIG.

7 is a sectional view of the burner of the present invention shown in FIG.

FIG. 8 is an explanatory view of a burner according to a second embodiment of the present invention viewed from a downstream direction.

[Explanation of symbols]

6 ... air, 8, 25 ... swirl vanes, 9 ... combustion chamber, 18 ... burner end face, 20 ... cooling air inlet hole, 21 ... cooling air passage, 22 ... fuel inlet hole, 24 ... fuel ejection hole, 33 ... premix Fuel nozzle, 34: cooling mixture inlet, 35: cooling mixture passage, 36: diffusion fuel inlet, 38: diffusion fuel ejection hole, 45: diffusion fuel, 46: premixed fuel.

──────────────────────────────────────────────────続 き Continued on the front page (51) Int.Cl. ⁶ Identification code Agency reference number FI Technical display location F23R 3/14 F23R 3/14 3/20 3/20 3/34 3/34

Claims (4)

[Claims] 1. A passage, which is located at an upstream end of a combustion chamber of a gas turbine and guides combustion air to the combustion chamber, a first swirler vane located in the passage, which gives a swirl to the combustion air.
In the gas turbine combustor provided with the first swirl vane portion, or a fuel nozzle that ejects a fuel gas into the passage downstream of the first swirl blade, a part of the combustion air is extracted. Guide to the cooling space provided at the burner end face,
Further, the gas turbine combustor has a structure in which the gas is discharged to the passage downstream of the first swirler through a second swirler provided in an outer peripheral portion of the cooling space. 2. The passage downstream of the first swirl vane according to claim 1, wherein a part of the combustion air is extracted and guided to a burner end face, and is communicated through a plurality of radially extending flow passages. A gas turbine combustor with a structure that allows the gas to flow out. 3. The gas turbine combustion chamber is located at an upstream end,
A first fuel nozzle for directly injecting combustion into the combustion chamber, a passage located on the outer periphery of the first fuel nozzle for guiding combustion air to the combustion chamber, and providing a swirl to the combustion air located in the passage A gas turbine device comprising: a first swirl vane; and a second fuel nozzle generally configured to inject fuel into the combustion air positioned in the passage and flowing through the passage to generate a mixture of fuel and air. A part of the air-fuel mixture is extracted and guided to the cooling space provided at the burner end face, and further discharged to the passage downstream of the first swirler through the second swirler provided at the outer periphery of the cooling space. A gas turbine combustor characterized by having a similar structure. 4. The method according to claim 3, wherein a part of the air-fuel mixture is bled and led to a burner end face, and the air-fuel mixture is provided downstream of the first swirl vane through a plurality of communicating radially extending flow paths. A gas turbine combustor with a structure that allows it to flow out into the passage.