JP2023153214A - Fuel nozzle and gas turbine combustor - Google Patents

Abstract

To provide a fuel nozzle comprising a plurality of fuel systems, which has less heat stress due to temperature difference of conducting fuel and combustion air, and a gas turbine combustor using the fuel nozzle.SOLUTION: A fuel nozzle comprises a plurality of flow paths, including a first flow path through which fuel or combustion air is conducted, and a second flow path through which fuel or combustion air is conducted and which is different from the first flow path. In the entire range where both the first flow path and the second flow path are arranged in a cross section perpendicular to an axial direction of the fuel nozzle, both the first flow path and the second flow path are divided into a plurality of parts and arranged in a circumferential direction of the fuel nozzle.SELECTED DRAWING: Figure 3

Description

The present invention relates to the structure of a fuel nozzle used in a gas turbine combustor, and particularly relates to a technique that is effective when applied to a pilot nozzle.

There are various types of fuel used in gas turbines, and the combustor to be used is selected depending on the fuel calorie and combustion speed. For low calorie fuels, a diffusion combustor is used, and for high calorie fuels, a premix combustor is used. Since premix combustion can reduce the flame temperature compared to diffusion combustion, it is possible to reduce NOx without spraying water or steam, and is currently widely applied to gas turbines.

Gas turbines used for power generation mainly use natural gas as fuel, and many natural gas-fired premix combustors are equipped with a pilot nozzle and a main nozzle, and the flame formed by the pilot nozzle is used to generate the main premix combustor. Efforts are being made to stabilize the mixed flame.

As background technology in this technical field, there is a technology such as that disclosed in Patent Document 1, for example. Patent Document 1 describes ``a pilot combustion burner for a gas turbine disposed at the axis of a combustor of a gas turbine, which includes a plurality of premix combustion fuel flow paths therein along the axial direction; a pilot combustion nozzle in which a plurality of fuel channels for diffusive combustion are formed independently; and a pilot combustion nozzle that is concentric with the pilot combustion nozzle and whose upstream end is downstream of the pilot combustion nozzle. a pilot burner tube arranged to surround an end of the pilot combustion nozzle; A gas turbine equipped with a plurality of swirling blades that apply swirling force to the compressed air passing through a ring-shaped air passage formed between the side ends and turn the compressed air into a swirling air flow. A pilot combustion burner is disclosed.

Japanese Patent Application Publication No. 2010-249449

As mentioned above, many natural gas-fired premix combustors are equipped with one pilot nozzle and eight main nozzles, and the fuel system consists of two systems: the main system and the pilot system. The pilot ratio (pilot fuel flow rate/total fuel flow rate) is highest at the time of ignition and decreases as the load increases, and the pilot ratio is lowest at rated load to suppress NOx emissions.

Additionally, if the methane concentration in the fuel changes, the combustibility will change, so adjust the fuel-air ratio in the combustion area by adjusting the air bypass valve, or change the pilot ratio to achieve a stable combustion state. This is necessary.

Incidentally, in the fuel nozzle of a gas turbine combustor, the occurrence of thermal stress due to the temperature difference between combustion air and fuel often poses a problem. If excessive thermal stress occurs, low cycle fatigue life will be insufficient and operational limitations will occur. In particular, in a fuel nozzle equipped with multiple fuel systems, such as the above-mentioned natural gas-fired premix combustor, fluids with different temperatures, such as fuel and combustion air (purge air), are conducted depending on the operating conditions. Thermal stress may increase due to
Thermal stress generated in the fuel nozzle leads to reduced reliability and durability of the fuel nozzle.

According to the above-mentioned Patent Document 1, it is possible to reduce vibrations caused by the flow of compressed air and prevent blow-off at the time of startup. ), no consideration is given to the thermal stress in the fuel nozzle that occurs due to the conduction of fluids with different temperatures.

SUMMARY OF THE INVENTION Therefore, an object of the present invention is to provide a fuel nozzle with a plurality of fuel systems that has less thermal stress caused by the temperature difference between the conducting fuel and combustion air, and a gas turbine combustor using the same. be.

In order to solve the above problems, the present invention provides a fuel nozzle including a plurality of flow paths, the first flow path through which fuel or combustion air flows, and the first flow path through which fuel or combustion air flows through. a second flow path different from the flow path, and in the entire range where both the first flow path and the second flow path are arranged in a cross section perpendicular to the axial direction of the fuel nozzle, The fuel nozzle is characterized in that both the first flow path and the second flow path are divided into a plurality of parts and arranged in a circumferential direction of the fuel nozzle.

The present invention also provides a combustor liner that constitutes a combustion chamber that combusts a mixture of fuel and combustion air, a transition piece that guides combustion gas from the combustion chamber to a turbine, and a transition piece that supplies fuel and combustion air to the combustion chamber. and a plurality of main nozzles arranged around the pilot nozzle to supply fuel and combustion air to the combustion chamber, the pilot nozzle having a first flow path through which fuel or combustion air flows. and a second flow path different from the first flow path through which fuel or combustion air is conducted, the first flow path and the second flow path being different from the first flow path in a cross section perpendicular to the axial direction of the pilot nozzle. The first flow path and the second flow path are both divided into a plurality of parts in the circumferential direction of the pilot nozzle and arranged in the entire range where both of the second flow paths are arranged. do.

According to the present invention, it is possible to realize a fuel nozzle having a plurality of fuel systems, which has less thermal stress caused by a temperature difference between conducting fuel and combustion air, and a gas turbine combustor using the fuel nozzle.

This makes it possible to provide a high-performance gas turbine combustor with excellent reliability and durability.

Problems, configurations, and effects other than those described above will be made clear by the following description of the embodiments.

FIG. 1 is a diagram showing an example of the configuration of a general gas turbine. FIG. 1 is a diagram showing an example of the configuration of a general combustor. 1 is a cross-sectional view showing the structure of a fuel nozzle according to Example 1 of the present invention. 4 is a sectional view taken along line A-A' in FIG. 3. FIG. 4 is a sectional view taken along line B-B' in FIG. 3. FIG. FIG. 2 is a cross-sectional view showing the structure of a conventional fuel nozzle. 6 is a sectional view taken along the line C-C' in FIG. 5. FIG. 6 is a sectional view taken along the line DD' in FIG. 5. FIG.

Embodiments of the present invention will be described below with reference to the drawings. Note that in each drawing, the same components are denoted by the same reference numerals, and detailed explanations of overlapping parts will be omitted.

First, with reference to FIGS. 1, 2, and 5 to 6B, a gas turbine combustor to which the present invention is applied and conventional problems will be described. FIG. 1 is a diagram showing an example of the configuration of a general gas turbine. FIG. 2 is a diagram showing an example of the configuration of a general combustor, and is shown as a combustor including a combustor liner 4 and a transition piece 5 that constitute a combustion chamber 15. FIG. 5 is a cross-sectional view showing the structure of a conventional pilot nozzle 7, and FIGS. 6A and 6B show a CC' cross section and a D-D' cross section in FIG. 5, respectively.

As shown in FIG. 1, a gas turbine is broadly divided into a compressor 1, a combustor 2, and a turbine 3. The compressor 1 adiabatically compresses air sucked in from the atmosphere as a working fluid, and the combustor 2 mixes fuel with the compressed air supplied from the compressor 1 and combusts it to generate high-temperature, high-pressure combustion gas. In this case, rotational power is generated when the combustion gas introduced from the combustor 2 expands. Exhaust gas from turbine 3 is released into the atmosphere.

As shown in FIG. 2, the combustor 2 includes a combustor lie 4 that constitutes a combustion chamber 15 that combusts a mixture of fuel and combustion air, and a combustor lie 4 that guides combustion gas from the combustion chamber 15 to a turbine 3 (combustion gas flow Direction 8) It includes a transition piece 5, a main nozzle 6 and a pilot nozzle 7 that supply fuel and combustion air to the combustion chamber 15. As described above, a plurality of main nozzles 6 (for example, eight) are arranged around one pilot nozzle 7.

As shown in FIG. 5, the conventional pilot nozzle 7 is constructed by joining together nozzle component members 9, 10, 11 in which flow passages A13 and flow passages B14 are formed in advance by drilling or the like at a joint portion 12. It is configured. For example, welding by brazing is used to join the nozzle constituent members 9, 10, and 11.

Generally, when the gas turbine is under rated load, relatively high-temperature sweep air (combustion air) is conducted through the flow path A13, and relatively low-temperature fuel such as natural gas is conducted through the flow path B14. Therefore, thermal stress is generated mainly due to the temperature difference in the radial direction of the pilot nozzle 7 and the resulting difference in thermal expansion in the radial direction and the axial direction. Generally, in a welded part, thermal stress is likely to be exacerbated due to the discontinuity of the shape due to unwelded parts, etc., and the fatigue strength of the welded part is lower than that of the base metal.

As described above, in the conventional pilot nozzle 7, the temperature difference between the fuel or combustion air that conducts the flow path A13 and the flow path B14, respectively, is This causes a strength bottleneck, and low cycle fatigue limits operation.

Further, as shown in FIG. 6A, in the root portion of the conventional pilot nozzle 7, both the flow path A13 and the flow path B14 are arranged in an annular shape in the circumferential direction of the pilot nozzle 7. , the pilot nozzle 7 has a structure that is thermally separated by a flow path A13 and a flow path B14 in the radial direction. Therefore, the thermal stress on the pilot nozzle 7 due to the temperature difference between the fuel or combustion air flowing through the flow path A13 and the flow path B14 is further aggravated.

Further, as shown in FIG. 6B, near the tip of the conventional pilot nozzle 7, the flow path B14 is divided into a plurality of parts in the circumferential direction of the pilot nozzle 7, but the flow path A13 is similar to the root part. The pilot nozzle 7 is arranged in an annular shape in the circumferential direction of the pilot nozzle 7, and the pilot nozzle 7 has a structure that is thermally separated by the flow path A13 in the radial direction.

Next, a fuel nozzle according to Example 1 of the present invention will be described with reference to FIGS. 3 to 4B. FIG. 3 is a cross-sectional view showing the structure of the pilot nozzle 7 of this embodiment, and FIGS. 4A and 4B show the A-A' cross-section and the B-B' cross-section of FIG. 3, respectively.

As shown in FIG. 3, the pilot nozzle 7 of this embodiment has a flow path A13 (first flow path) through which fuel or combustion air flows and a flow path A13 (first flow path) through which fuel or combustion air conducts. It has a flow path B14 (second flow path) different from the flow path), and among the nozzle constituent members 9 and 10 of the pilot nozzle 7, at least the flow path A13 (first flow path) and the flow path B14 (Second flow path) The portion where both of the two flow paths are arranged is constituted by an integral nozzle component 10 without a joint portion 12.

As shown in FIG. 3, the part where both the flow path A13 (first flow path) and the flow path B14 (second flow path) are arranged is configured with an integral nozzle component 10 without a joint part 12. Therefore, it is possible to prevent the joint 12 from becoming a strength bottleneck due to the temperature difference between the fuel or combustion air that conducts the flow path A13 and the flow path B14, as described above, and the pilot nozzle 7 The reliability and durability of the system can be improved.

Further, as shown in FIGS. 4A and 4B, in the pilot nozzle 7 of this embodiment, the flow path A13 (first flow path) and the flow path B14 (second flow path) are arranged in the circumferential direction of the pilot nozzle 7. Both are divided into multiple parts and arranged.

As shown in FIGS. 4A and 4B, both the flow path A13 (first flow path) and the flow path B14 (second flow path) are divided into a plurality of parts and arranged in the circumferential direction of the pilot nozzle 7. This prevents the pilot nozzle 7 from being completely thermally separated by the flow path A13 (first flow path) and flow path B14 (second flow path) in the radial direction. Thereby, thermal stress on the pilot nozzle 7 due to the temperature difference between the fuel or combustion air flowing through the flow path A13 and the flow path B14 can be alleviated.

For example, even if combustion air is conducted through the flow path A13 (first flow path) and fuel whose temperature is lower than that of the combustion air is conducted through the flow path B14 (second flow path), the fuel and Since the thermal stress on the pilot nozzle 7 due to the temperature difference of the combustion air is alleviated, in addition to the effect of configuring the nozzle component 10 as a single piece without a joint 12, the reliability and durability of the pilot nozzle 7 are improved. Further improvements can be made.

Further, as shown in FIG. 3, the nozzle component 9 near the tip of the pilot nozzle 7 has a flow path A13 (first flow path) and a flow path B14 (second flow path). The nozzle component 9 has only the flow path A13 (first flow path) and the flow path B14 (first flow path). The second flow path) is joined to the nozzle component 10 in which both are arranged, for example, by welding by brazing or by hot isostatic pressing (HIP).

As shown in FIG. 3, in a region where only one of the flow paths A13 (first flow path) and flow path B14 (second flow path) is formed, By arranging the joints 12 in a limited manner, it is possible to prevent thermal stress from occurring on the pilot nozzle 7 due to a temperature difference between the fuel or combustion air flowing through the flow path, and the joints 12 can be arranged in a limited manner. The bonding reliability can be guaranteed.

Note that it is desirable to use the above-mentioned hot isostatic pressing (HIP) method for joining the nozzle component 9 and the nozzle component 10 at the joint 12. By using the hot isostatic pressing (HIP) method, unwelded parts can be eliminated as much as possible, so thermal stress caused by shape discontinuity in the joint part 12 can be suppressed.

As explained above, according to the present invention, it is possible to realize a fuel nozzle with less thermal stress caused by the temperature difference between conducting fuel and combustion air, and a gas turbine combustor using the same. The reliability and durability of the system can be improved.

Note that the present invention is not limited to the above-described embodiments, and includes various modifications.
For example, the embodiments described above are described in detail to explain the present invention in an easy-to-understand manner, and the present invention is not necessarily limited to having all the configurations described. Furthermore, it is possible to replace a part of the configuration of one embodiment with the configuration of another embodiment, and it is also possible to add the configuration of another embodiment to the configuration of one embodiment. Further, it is possible to add, delete, or replace a part of the configuration of each embodiment with other configurations.

1... Compressor 2... Combustor 3... Turbine 4... Combustor liner 5... Transition piece (transition piece)
6... Main nozzle 7... Pilot nozzle 8... Flow direction of combustion gas 9, 10, 11... Nozzle constituent members 12... Joint portion 13... Flow path A
14...Flow path B
15...Combustion chamber

Claims (10)

A fuel nozzle including a plurality of flow paths, the fuel nozzle comprising:
a first flow path through which fuel or combustion air is conducted;
a second flow path different from the first flow path through which fuel or combustion air is conducted;
In the entire range where both the first flow path and the second flow path are arranged in a cross section perpendicular to the axial direction of the fuel nozzle, both the first flow path and the second flow path are arranged. A fuel nozzle, characterized in that the fuel nozzle is divided into a plurality of parts and arranged in a circumferential direction. The fuel nozzle according to claim 1,
The fuel nozzle is configured such that at least a region in a cross section perpendicular to the axial direction of the fuel nozzle in which both the first flow path and the second flow path are arranged is an integral member with no joints. A fuel nozzle characterized by: The fuel nozzle according to claim 1 or 2,
Combustion air is conducted through the first flow path,
A fuel nozzle, characterized in that fuel having a lower temperature than the combustion air flows through the second flow path. The fuel nozzle according to any one of claims 1 to 3,
Of the first flow path and the second flow path, only the first flow path is arranged near the tip of the fuel nozzle,
A fuel nozzle characterized in that a portion where only the first flow path is arranged is joined to a portion where the first flow path and the second flow path are arranged. The fuel nozzle according to claim 4,
The region where only the first channel is arranged is separated from the region where the first channel and the second channel are arranged by welding or hot isostatic pressing (HIP). A fuel nozzle characterized in that it is joined. a combustor liner that constitutes a combustion chamber that combusts a mixture of fuel and combustion air;
a transition piece that guides combustion gas from the combustion chamber to the turbine;
a pilot nozzle that supplies fuel and combustion air to the combustion chamber;
A plurality of main nozzles are arranged around the pilot nozzle and supply fuel and combustion air to the combustion chamber,
The pilot nozzle includes a first flow path through which fuel or combustion air is conducted;
a second flow path different from the first flow path through which fuel or combustion air is conducted;
In the entire range where both the first flow path and the second flow path are arranged in a cross section perpendicular to the axial direction of the pilot nozzle, both the first flow path and the second flow path are arranged. A gas turbine combustor characterized in that the pilot nozzle is divided into a plurality of parts and arranged in a circumferential direction. The gas turbine combustor according to claim 6,
The pilot nozzle is configured such that at least a region in a cross section perpendicular to the axial direction of the pilot nozzle in which both the first flow path and the second flow path are arranged is an integral member with no joints. A gas turbine combustor characterized by: The gas turbine combustor according to claim 6 or 7,
Combustion air is conducted through the first flow path,
A gas turbine combustor, wherein fuel having a temperature lower than that of the combustion air flows through the second flow path. The gas turbine combustor according to any one of claims 6 to 8,
Of the first flow path and the second flow path, only the first flow path is arranged near the tip of the pilot nozzle,
A gas turbine combustor characterized in that a portion where only the first flow path is arranged is joined to a portion where the first flow path and the second flow path are arranged. The gas turbine combustor according to claim 9,
The region where only the first channel is arranged is separated from the region where the first channel and the second channel are arranged by welding or hot isostatic pressing (HIP). A gas turbine combustor characterized by being joined.

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