Combustor Fuel Nozzle Structure

Abstract

A fuel nozzle including a plurality of fuel lines is provided, which has low thermal stress caused by a temperature difference between fuel and combustion air which are passed through the fuel nozzle, and also a gas turbine combustor using the fuel nozzle is provided. The fuel nozzle includes a plurality of channels including a first channel through which either fuel or combustion air is passed, and a second channel through which either fuel or combustion air is passed and which is distinct from the first channel. Of components of the fuel nozzle, a single-piece component makes up at least a region where the first channel and the second channel are placed.

Description

CLAIM OF PRIORITY

The present application claims priority from Japanese Patent application serial no. 2020-155193, filed on Sep. 16, 2020, the content of which is hereby incorporated by reference into this application.

BACKGROUND OF THE INVENTION

The present invention relates to structure of a fuel nozzle used in a gas turbine combustor and, more particularly, to an effective technique in application to a pilot nozzle.

There are various types of fuel for use in gas turbines, and an appropriate combustor is selected based on fuel calorie and burning rate. Low-calorific fuel is appropriate for use in a diffusion combustor, and high-calorific fuel is appropriate for use in a premix combustor. Premix combustion provides a reduction in flame temperature as compared with diffusion combustion. Therefore, the premix combustion can achieve reduced NOx without a spray of water or steam, and is widely applied to gas turbines today.

In gas turbines used for electric power generation, natural gas is mainly used as fuel. Many natural gas burning premix combustors include a pilot nozzle and main nozzles and seek stabilization of main premixed flame by flame formed by the pilot nozzle.

As one of conventional techniques in such technological field, for example, Japanese Unexamined Patent Application Publication No. 2010-249449 discloses as follows. “A gas turbine pilot combustion burner placed at an axis of a combustor of a gas turbine, comprising: a pilot combustion nozzle having a plurality of premix combustion fuel channels and a plurality of diffusion combustion fuel channels formed independently therein along an axial direction; a pilot burner cylinder that is disposed concentrically with respect to the pilot combustion nozzle and in conditions where an upstream end of the pilot burner cylinder surrounds a downstream end of the pilot combustion nozzle; and a plurality of swirl vanes that are radially disposed on the downstream end of the pilot combustion nozzle to apply swirl force to compressed air passing through a ring-shaped air passage in order to convert the compressed air to a swirl air flow, the ring-shaped air passage being formed between the downstream end of the pilot combustion nozzle and the upstream end of the burner cylinder.”

As described above, many natural gas burning premix combustors include a pilot nozzle and eight main nozzles, and the fuel lines mainly includes two lines, a main line and a pilot line. A pilot ratio (pilot fuel flow rate/total fuel flow rate) is highest at ignition, and then decreases with increase in load. And, at rated load, the pilot ratio is minimized for a reduction in NOx emission.

Also, as methane concentration in fuel changes, the combustion characteristics change. This involves adjusting an air bypass valve to adjust a fuel-air ratio in the combustion region, and/or changing a pilot ratio for adjustment for a stable combustion state.

In this connection, the fuel nozzle of the gas turbine combustor often has a problem of producing thermal stress caused by a temperature difference between the combustion air and the fuel. Excessive thermal stress causes inadequate life of low cycle fatigue, resulting in restriction of operation. In particular, in the fuel nozzle including a plurality of fuel lines such as in the aforementioned natural gas burning premix combustor, several fluids with different temperatures, such as fuel, combustion air (purge air), and the like, are passed through the fuel nozzle depending on operating conditions, and this may cause an increase in thermal stress. The thermal stress induced in the fuel nozzle leads to a reduction in reliability and durability of the fuel nozzle.

According to Japanese Unexamined Patent Application Publication No. 2010-249449, vibrations produced by the compressed air flowing are mitigated, and also blowing-off at startup is prevented. However, no consideration is given to the thermal stress caused on the fuel nozzle by passage of the fluids with different temperatures such as fuel, combustion air (purge air) and the like, as described above.

SUMMARY OF THE INVENTION

Accordingly, it is an object of the present invention to provide a fuel nozzle including a plurality of fuel lines with low thermal stress caused by a temperature difference between fuel and combustion air which are passed through the fuel nozzle, and also to provide a gas turbine combustor using the fuel nozzle.

To achieve the above object, in an aspect of the present invention, a fuel nozzle includes: a plurality of channels including a first channel through which either fuel or combustion air is passed; and a second channel through which either fuel or combustion air is passed and which is distinct from the first channel. Of components of the fuel nozzle, a single-piece component of the components of the fuel nozzle makes up at least a region where the first channel and the second channel are placed.

Further, in another aspect of the present invention, a gas turbine combustor includes: a combustor liner that essentially makes up a combustion chamber in which a gas mixture of fuel and combustion air is burned; a transition piece through which combustion gases are directed from the combustion chamber toward a turbine; a pilot nozzle that supplies the fuel and the combustion air into the combustion chamber; and a plurality of main nozzles that are arranged around the pilot nozzle to supply the fuel and the combustion air into the combustion chamber. The pilot nozzle has: a first channel through which either the fuel or the combustion air is passed; and a second channel through which either the fuel or the combustion air is passed and which is distinct from the first channel. The pilot nozzle includes components, and a single-piece component of the components of the pilot nozzle makes up at least a region where the first channel and the second channel are placed.

According to the present invention, it is possible to implement a fuel nozzle which includes a plurality of fuel lines and has low thermal stress caused by a temperature difference between fuel and combustion air which are passed through the fuel nozzle, and also a gas turbine combustor using the fuel nozzle.

This enables the high-performance gas turbine combustor excelling in reliability and durability.

These and other objects, features and advantages will be apparent from a reading of the following description of embodiments.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a diagram illustrating example configuration of a typical gas turbine;

FIG. 2 is a diagram illustrating example configuration of a typical combustor;

FIG. 3 is a cross-sectional view illustrating fuel nozzle structure in Example 1 according to the present invention;

FIG. 4A is a view of cross section A-A′ in FIG. 3;

FIG. 4B is a view of cross section B-B′ in FIG. 3;

FIG. 5 is a cross-sectional view illustrating conventional fuel nozzle structure;

FIG. 6A is a view of cross section C-C′ in FIG. 5; and

FIG. 6B is a view of cross section D-D′ in FIG. 5.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

Examples according to the present invention will now be described with reference to the accompanying drawings. It is to be understood that like reference signs indicate the same or similar configurations throughout the drawings, and a detailed description of duplicated portion is omitted.

Example 1

First, a gas turbine combustor according to the present invention, and conventional problems are described with reference to FIGS. 1 and 2 and FIGS. 5 to 6B. FIG. 1 is a diagram illustrating example configuration of a typical gas turbine. FIG. 2 is a diagram illustrating example configuration of a typical combustor, illustrated as a combustor including a combustor liner 4 essentially making up a combustion chamber 15, and a transition piece 5. FIG. 5 is a cross-sectional view illustrating conventional structure of a pilot nozzle 7. FIGS. 6A and 6B are views of cross section C-C′ and cross section D-D′ in FIG. 5, respectively.

As illustrated in FIG. 1, a gas turbine is roughly made up of compressor 1, a combustor 2, and a turbine 3. The compressor 1 adiabatically compresses, as working fluid, air taken in from the atmosphere. The combustor 2 mixes and burns fuel with the compressed air fed from the compressor 1 in order to generate high-temperature, high-pressure combustion gases. Then, in the turbine 3, rotative power is produced when expanding the combustion gases introduced from the combustor 2. The emission from the turbine 3 is released into the atmosphere.

As illustrated in FIG. 2, the combustor 2 includes: a combustor liner 4 essentially making up a combustion chamber 15 in which a gas mixture of fuel and combustion air is burned; a transition piece 5 through which the combustion gases are directed from the combustion chamber 15 toward the turbine 3 (the flow direction 8 of combustion gases); and main nozzles 6 and a pilot nozzle 7 which supply fuel and combustion air into the combustion chamber 15. The plurality of main nozzles 6 (e.g., eight main nozzles 6) are arranged around the single pilot nozzle 7 as described earlier.

As illustrated in FIG. 5, the conventional pilot nozzle 7 is configured such that nozzle components 9, 10, 11 are joined together at joints 12, the nozzle components 9, 10, 11 having a channel A13 and a channel B14 pre-formed therein through hole machining such as drilling. The nozzle components 9, 10, 11 are joined together by use of, for example, braze welding.

Typically, under the rated load of the gas turbine, relatively high temperature sweep air (combustion air) is passed through the channel A13, and fuel, such as relatively low temperature natural gas, is passed through the channel B14. Therefore, thermal stress occurs due to a temperature difference mainly in the radial direction of the pilot nozzle 7 and a thermal expansion difference in the radial direction and the axial direction which arises from the temperature difference. Typically, in the welded area, shape discontinuity due to an unwelded portion and/or the like causes the thermal stress to be readily promoted and the fatigue strength in the welded area is reduced as compared with that in the base material.

Hence, in the conventional pilot nozzle 7, in particular, the joint 12 corresponding to a region where both the channel A13 and the channel B14 are placed becomes a severe bottleneck due to the temperature difference of the fuel or the combustion air which are passed individually through the channel A13 or the channel B14, and thus the operation is restricted by low cycle fatigue.

As illustrated in FIG. 6A, because in a base portion of the conventional pilot nozzle 7, both the channel A13 and the channel B14 are disposed annularly in the circumferential direction of the pilot nozzle 7, the pilot nozzle 7 is structured to be thermally divided in the radial direction by the channel A13 and the channel B14. Therefore, the thermal stress on the pilot nozzle 7 is further promoted by the temperature difference of the fuel or the combustion air which are passed independently through the channel A13 and the channel B14.

Also, as illustrated in FIG. 6B, in proximity to the distal end of the conventional pilot nozzle 7, the channel B14 is divided into a plurality of sections and arranged in the circumferential direction of the pilot nozzle 7, while, as in the base portion, the channel A13 is disposed annularly in the circumferential direction of the pilot nozzle 7. Thus, the pilot nozzle 7 is structured to be thermally divided in the radial direction by the channel A13.

A fuel nozzle in Example 1 according to the present invention will now be described with reference to FIG. 3 to FIG. 4B. FIG. 3 is a cross-sectional view illustrating the structure of the pilot nozzle 7 in Example 1. FIGS. 4A and 4B are views of cross section A-A′ and cross section B-B′ in FIG. 3.

As illustrated in FIG. 3, the pilot nozzle 7 in Example 1 has a channel A13 (first channel) through which fuel or combustion air is passed, and a channel B14 (second channel) through which fuel or combustion air is passed and which is distinct from the channel A13 (first channel). Of the nozzle components 9, 10 of the pilot nozzle 7, the nozzle component 10 which is a single piece without the joint 12 makes up at least a region where both the channel A13 (first channel) and the channel B14 (second channel) are placed.

As illustrated in FIG. 3, the region where both the channel A13 (first channel) and the channel B14 (second channel) are placed is made up of the single-piece nozzle component 10 without the joint 12. This enables prevention of the joint 12 from acting as a severe bottleneck due to a temperature difference of the fuel or the combustion air which are passed individually through the channel A13 and the channel B14, as described earlier. Thus, it is possible to improve the reliability and durability of the pilot nozzle 7.

As illustrated in FIGS. 4A and 4B, regarding the pilot nozzle 7 in the example, each of the channel A13 (first channel) and the channel B14 (second channel) is divided into a plurality of sections and arranged in the circumferential direction of the pilot nozzle 7.

As illustrated in FIGS. 4A and 4B, each of the channel A13 (first channel) and the channel B14 (second channel) is divided into a plurality of sections and arranged in the circumferential direction of the pilot nozzle 7, whereby the pilot nozzle 7 is prevented from being completely, thermally divided in the radial direction by the channel A13 (first channel) and the channel B14 (second channel). In turn, this makes it possible to relieve the thermal stress on the pilot nozzle 7 due to the temperature difference of the fuel or the combustion air which are passed individually through the channel A13 and the channel B14.

For example, even where the combustion air is passed through the channel A13 (first channel) and the fuel with a lower temperature than the combustion air is passed through the channel B14 (second channel), the thermal stress on the pilot nozzle 7 due to a temperature difference between the fuel and the combustion air is relieved. Because of this, in addition to the effect of configuration using the single-piece nozzle component 10 without the joint 12, further improvement in reliability and durability of the pilot nozzle 7 is enabled.

Also, as illustrated in FIG. 3, of both the channel A13 (first channel) and the channel B14 (second channel), only the channel A13 (first channel) is disposed in the nozzle component 9 in proximity to the distal end of the pilot nozzle 7. The nozzle component 9 with only the channel A13 (first channel) disposed therein is joined to the nozzle component 10 with both the channel A13 (first channel) and the channel B14 (second channel) disposed therein, for example, by braze welding or HIP (Hot Isostatic Pressing) technique.

As illustrated in FIG. 3, the joint 12 is disposed exclusively in the area in which only one channel A13 (first channel) of both the channel A13 (the first channel) and the channel B14 (the second channel) is formed. This structure enables prevention of the thermal stress from being induced on the pilot nozzle 7 by a temperature difference of the fuel or combustion air passed through the channel in question, and in turn ensuring joint reliability of the joint 12.

It is noted that the above-described HIP (Hot Isostatic Pressing) technique is desirably used to join the nozzle component 9 and the nozzle component 10 together at the joint 12. Using the HIP technique enables maximum elimination of unwelded area. This enables suppression of the thermal stress caused by shape discontinuity in the joint 12.

As described above, according to the present invention, it is possible to provide a fuel nozzle with low thermal stress caused by a temperature difference between fuel and combustion air which are passed therethrough, and a gas turbine combustor using the fuel nozzle, and improvements in the reliability and durability of the gas turbine combustor can be made.

It should be understood that the present invention is not limited to the above examples and is intended to embrace various modifications. For example, the above examples have been described in detail for the purpose of explaining the present invention clearly, and the present invention is not necessarily limited to including all the components and configurations described above. Further, a portion of the configuration in one example may be substituted for configuration in another example and configuration in one example may be added to configuration in another example. Further, on a portion of the configuration in each example, addition, deletion and substitution of another configuration may be made.

REFERENCE SIGNS LIST

• 1 . . . Compressor
• 2 . . . Combustor
• 3 . . . Turbine
• 4 . . . Combustor liner
• 5 . . . Transition piece
• 6 . . . Main nozzle
• 7 . . . Pilot nozzle
• 8 . . . Flow direction of combustion gases
• 9, 10, 11 . . . Nozzle component
• 12 . . . Joint
• 13 . . . Channel A
• 14 . . . Channel B
• 15 . . . Combustion chamber

Claims

1. A fuel nozzle, comprising a plurality of channels including: a first channel through which either fuel or combustion air is passed; anda second channel through which either fuel or combustion air is passed and which is distinct from the first channel,wherein the fuel nozzle includes components, and a single-piece component of the components of the fuel nozzle makes up at least a region where the first channel and the second channel are placed.

2. The fuel nozzle according to claim 1, wherein each of the first channel and the second channel is divided into a plurality of sections and arranged in a circumferential direction of the fuel nozzle.

3. The fuel nozzle according to claim 1, wherein the combustion air is passed through the first channel, andthe fuel of a lower temperature than that of the combustion air is passed through the second channel.

4. The fuel nozzle according to claim 1, wherein, of the first channel and the second channel, only the first channel is disposed in proximity to a distal end of the fuel nozzle, anda region where only the first channel is placed is joined to the region where the first channel and the second channel are placed.

5. The fuel nozzle according to claim 4, wherein the region where only the first channel is placed is joined to the region where the first channel and the second channel are placed, by either welding or Hot Isostatic Pressing (HIP) technique.

6. A gas turbine combustor, comprising: a combustor liner that essentially makes up a combustion chamber in which a gas mixture of fuel and combustion air is burned;a transition piece through which combustion gases are directed from the combustion chamber toward a turbine;a pilot nozzle that supplies the fuel and the combustion air into the combustion chamber; anda plurality of main nozzles that are arranged around the pilot nozzle to supply the fuel and the combustion air into the combustion chamber,wherein the pilot nozzle has: a first channel through which either the fuel or the combustion air is passed; anda second channel through which either the fuel or the combustion air is passed and which is distinct from the first channel, andthe pilot nozzle includes components, and a single-piece component of the components of the pilot nozzle makes up at least a region where the first channel and the second channel are placed.

7. The gas turbine combustor according to claim 6, wherein each of the first channel and the second channel is divided into a plurality of sections and arranged in a circumferential direction of the pilot nozzle.

8. The gas turbine combustor according to claim 6, wherein the combustion air is passed through the first channel, andthe fuel of a lower temperature than that of the combustion air is passed through the second channel.

9. The gas turbine combustor according to claim 6, wherein, of the first channel and the second channel, only the first channel is disposed in proximity to a distal end of the pilot nozzle, anda region where only the first channel is placed is joined to the region where the first channel and the second channel are placed.

10. The gas turbine combustor according to claim 9, wherein the region where only the first channel is placed is joined to the region where the first channel and the second channel are placed, by either welding or Hot Isostatic Pressing (HIP) technique.