RU2774929C1 - Fuel injector and gas turbine combustion chamber - Google Patents

Abstract

FIELD: gas turbine engines.

SUBSTANCE: invention relates to the combustion chambers of gas turbine engines. The fuel injector contains a plurality of channels and includes: a first channel through which fuel or combustion air passes, a second channel through which fuel or combustion air passes and which is different from the first channel, and the fuel injector includes structural elements, and an integral structural element of the fuel injector of these structural elements is at least the area in which the first channel and the second channel are located. The first channel and the second channel are divided into a plurality of sections arranged in the circumferential direction of the fuel injector.

EFFECT: invention makes it possible to obtain a fuel injector with low thermal stress caused by the temperature difference between the fuel and combustion air that pass through it, as well as to increase the reliability and service life of the gas turbine combustion chamber.

10 cl, 8 dwg

Description

Prerequisites for the creation of the invention

The present invention relates to the design of a fuel injector used in the combustion chamber of a gas turbine and, in particular, to an effective technical solution in application to a pilot injector.

There are various types of fuels for gas turbine applications and the appropriate combustor is selected depending on the calorific value of the fuel and the burning rate. Low calorie fuel is suitable for use in a diffusion combustor and high calorific fuel is suitable for use in a premix combustor. Premix combustion provides a reduction in flame temperature compared to diffusion combustion. Therefore, premix combustion can reduce NOx emissions without splashing water or steam, and is currently widely used in gas turbines.

Gas turbines used to generate electricity primarily use natural gas as fuel. Many natural gas premix combustors are equipped with a pilot burner and main burners and stabilize the premix main flame with the flame generated by the pilot burner.

As one of the prior art solutions in this field, for example, Japanese Patent Laid-Open Publication No. 2010-249449 discloses the following. "A pilot combustion burner in a gas turbine, located on the axis of the combustion chamber of a gas turbine, containing: a pilot combustion nozzle having a plurality of fuel channels for pre-mixing combustion and a plurality of fuel channels for diffusion combustion formed independently in the axial direction; a pilot burner cylinder , which is placed concentrically with respect to that combustion pilot nozzle such that the upstream end of the pilot burner cylinder surrounds the downstream end of the combustion pilot nozzle; air passing through an annular air passage formed between the downstream end of the combustion pilot nozzle and the upstream end of the burner cylinder to convert the compressed air into a swirl airflow."

As mentioned above, many natural gas premix combustors include one pilot injector and eight main injectors, and the fuel lines generally include two fuel lines, the main fuel line and the pilot fuel line. The pilot ratio (pilot fuel consumption/total fuel consumption) is highest at ignition and then decreases with increasing load, and at rated load the pilot ratio is lowest, resulting in NOx reduction.

In addition, when the concentration of methane in the fuel changes, the combustion characteristics also change. Therefore, it becomes necessary to control the fuel-air ratio in the combustion zone by means of an air bypass valve and/or change the pilot ratio in order to achieve a stable combustion state.

In this case, the problem of the fuel injector of the combustion chamber of a gas turbine is often the occurrence of thermal stress caused by a temperature difference between the combustion air and the fuel. Excessive thermal stress leads to reduced life due to low cycle fatigue and limited service capability. Specifically, in a fuel injector including a plurality of fuel lines, as in the above-mentioned natural gas premix combustion chamber, multiple fluids of different temperatures such as fuel and combustion air (scavenging air) and the like. , pass through the fuel injector, depending on the operating conditions, and this can lead to an increase in thermal stress. Thermal stress generated in the fuel injector leads to reduced reliability and shortened life of the fuel injector.

According to Japanese Patent Laid-Open Publication No. 2010-249449, the vibrations generated by the compressed air flow are reduced and the blowout at startup is also prevented. However, the thermal stress caused on the fuel injector by the passage of fluids of different temperatures such as the above-mentioned fuel, combustion air (scavenging air) and the like is not taken into account.

Brief summary of the invention

Therefore, it is an object of the present invention to provide a fuel injector including a plurality of fuel lines with low thermal stress caused by the temperature difference between the fuel and combustion air passing through the fuel injector, as well as the combustion chambers of a gas turbine using the fuel injector.

In order to achieve the above object, in an aspect of the present invention, the fuel injector includes a plurality of passages: including a first passage through which fuel or combustion air flows; and a second passage through which the fuel or combustion air passes and which is different from the first passage. Of the structural elements of the fuel injector, the one-piece structural element of the fuel injector is at least the area in which the first channel and the second channel are located.

Further, in another aspect of the present invention, a gas turbine combustor includes: a combustor liner that substantially constitutes a section of a combustor in which a mixture of fuel and combustion air is burned; a transition compartment through which combustion gases are directed from the combustion chamber section to the turbine; a pilot nozzle that supplies fuel and combustion air to the combustion chamber section; and a plurality of main injectors that are positioned around the pilot injector and supply fuel and combustion air to the combustion chamber section. The pilot nozzle has: a first passage through which fuel or combustion air passes; and a second passage through which the fuel or combustion air passes and which is different from the first passage. The pilot injector includes structural elements, and the one-piece pilot injector structural element of these structural elements constitutes at least the area in which the first channel and the second channel are located.

According to the present invention, it is possible to realize a fuel injector that includes a plurality of fuel lines and has a low thermal stress caused by a temperature difference between fuel and combustion air passing through the fuel injector, as well as a gas turbine combustion chamber using the fuel injector.

This results in a highly efficient gas turbine combustor with high reliability and long service life.

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

Brief description of the drawings

Fig. 1 is a schematic illustration of an embodiment of a typical gas turbine design;

Fig. 2 is a schematic illustration of an embodiment of a typical combustion chamber design;

Fig. 3 is a sectional view illustrating the construction of a fuel injector according to Embodiment 1 of the present invention;

Fig. 4A is a view of the fuel injector shown in FIG. 3, in section along A-A';

Fig. 4B is a view of the fuel injector shown in FIG. 3. in a section along В-В';

Fig. 5 is a sectional view illustrating the construction of a prior art fuel injector;

Fig. 6A is a view of the fuel injector shown in FIG. 5, in a section along C-C;

is a cross-sectional view in FIG. 5; and

Fig. 6B is a view of the fuel injector shown in FIG. 5, in section along D-D'.

Detailed Description of the Preferred Embodiments

Below with reference to the accompanying drawings is a description of embodiments of the present invention. In each drawing, however, the same or similar structural elements are designated by the same reference numerals, in the repetition of which their detailed description is not given.

Embodiment 1

First with reference to FIG. 1 and 2 and in Fig. 5-6B describe the combustion chamber of a gas turbine in accordance with the present invention and prior art problems. In FIG. 1 is a schematic illustration of an embodiment of a typical gas turbine design. In FIG. 2 is a schematic illustration of an embodiment of a typical combustion chamber design shown as a combustion chamber including a chamber insert 4 substantially constituting a combustion chamber section 15 and an adapter compartment 5. FIG. 5 is a sectional view illustrating the construction of a pilot nozzle 7 in the prior art. In FIG. 6A and FIG. 6B are views of the fuel injector shown in FIG. 5, sectioned along C-C' and D-D', respectively.

As shown in FIG. 1, a gas turbine generally consists of a compressor 1, a combustor 2 and a turbine 3. The compressor 1 adiabatically compresses the air sucked in from the atmosphere as a working fluid. Combustion chamber 2 mixes and burns fuel with compressed air supplied from compressor 1, resulting in combustion gases of high temperature and high pressure. With the subsequent expansion of the combustion gases coming from the combustion chamber 2, the turbine 3 generates a rotational force. The exhaust gases from the turbine 3 are released into the atmosphere.

As shown in FIG. 2, the combustion chamber 2 includes: a combustion chamber liner 4 essentially constituting a combustion chamber section 15 in which a mixture of fuel and combustion air is burned; a transition compartment 5 through which the combustion gases are directed from the section 15 of the combustion chamber towards the turbine 3 (in the direction 8 of the flow of combustion gases); as well as the main injectors 6 and the pilot injector 7 which supply fuel and combustion air to the combustion chamber section 15. As described above, a plurality of main nozzles 6 (for example, eight main nozzles 6) are placed around one pilot nozzle 7.

As shown in FIG. 5, the prior art pilot nozzle 7 has such a construction in which the nozzle elements 9, 10, 11 are connected to each other at the connection sections 12, while the nozzle elements 9, 10, 11 have a channel A13 and a channel B14 preformed in them as a result of mechanical processing, such as drilling. The nozzle elements 9, 10, 11 are connected to each other using, for example, solder welding.

Typically, at rated load of the gas turbine, purge air (combustion air) at a relatively high temperature passes through port A13 and fuel such as natural gas at a relatively low temperature passes through port B14. Therefore, due to the temperature difference mainly in the radial direction of the pilot nozzle 7 and the thermal expansion difference in the radial direction and the axial direction caused by this temperature difference, thermal stress is generated. As a rule, non-uniformity of the shape due to the appearance of non-welded areas and/or the like. leads to an increase in thermal stress and a decrease in fatigue strength in the welding area compared to the base material.

Therefore, in the prior art pilot nozzle 7, in particular, the connection section 12 corresponding to the area of the location of both the channel A13 and the channel B14 becomes a serious bottleneck due to the temperature difference of the fuel or combustion air, which pass through the channel A13 or the channel respectively. B14, and the performance of this injector is limited by low cycle fatigue.

As shown in FIG. 6A, since in the base of the prior art pilot nozzle 7, both the passage A13 and the passage B14 are arranged annularly in the circumferential direction of the pilot nozzle 7, the pilot nozzle 7 has a structure thermally separated in the radial direction by the passage A13 and the passage B14. Therefore, the thermal stress on the pilot nozzle 7 is further enhanced by the temperature difference of the fuel or combustion air, which pass through the passage A13 and the passage B14, respectively.

In addition, as shown in FIG. 6B, near the front end of the prior art pilot nozzle 7, channel B14 is divided into a plurality of channels located in the circumferential direction of the pilot nozzle 7, and channel A13 is arranged, as in the base, annularly in the circumferential direction of the pilot nozzle 7. Thus, the pilot the nozzle 7 has a structure thermally separated in the radial direction by the channel A13.

Below with reference to Fig. 3-4B is a description of a fuel injector according to Embodiment 1 of the present invention. In FIG. 3 is a sectional view illustrating the structure of the pilot nozzle 7 according to Embodiment 1 of the present invention. In FIG. 4A and 4B are views of the pilot nozzle shown in FIG. 3, in section along A-A' and B-B' respectively.

As shown in FIG. 3, the pilot nozzle 7 in Embodiment 1 has a port A13 (first port) through which the fuel or combustion air flows, and a port B14 (second port) through which the fuel or combustion air passes and which is different from port A13 (from first channel). Of the injector structural elements 9, 10 in the pilot injector 7, the structural element 10, which is an integral structural element without a connection section 12, constitutes at least an area in which both channel A13 (first channel) and channel B14 (second channel) are located.

As shown in FIG. 3, the area in which both the channel A13 (first channel) and the channel B14 (second channel) are located, consists of an integral structural element 10 without a connection section 12. This prevents connection section 12 from becoming a serious bottleneck due to, as described above, the temperature difference of the fuel or combustion air that passes through channel A13 and channel B14, respectively. Thus, it is possible to improve the reliability and increase the service life of the pilot nozzle 7.

As shown in FIG. 4A and 4B relating to the pilot nozzle 7 in the embodiment, and channel A13 (first channel) and channel B14 (second channel) are divided into a plurality of channels disposed in the circumferential direction of the pilot nozzle 7.

As shown in FIG. 4A and 4B, both passage A13 (first passage) and passage B14 (second passage) are divided into a plurality of passages arranged in the circumferential direction of the pilot nozzle 7, thereby preventing complete thermal separation of the pilot nozzle 7 in the radial direction by passage A13 (first passage ) and channel B14 (second channel). In turn, this makes it possible to reduce the thermal stress on the pilot nozzle 7 due to the temperature difference of the fuel or combustion air, which pass through the channel A13 and channel B14, respectively.

For example, even in the case where the combustion air passes through passage A13 (the first passage) and fuel with a lower temperature than the combustion air passes through passage B14 (the second passage), the thermal stress on the pilot injector 7 due to the temperature difference between fuel and combustion air is reduced. Therefore, in addition to the design effect of using a one-piece nozzle structure member 10 without a connection section 12, an additional increase in reliability and an increase in the service life of the pilot nozzle 7 is provided.

In addition, as shown in FIG. 3, from both channels - channel A13 (first channel) and channel B14 (second channel) - only channel A13 (first channel) is located in the element 9 of the nozzle structure near the front end of the pilot nozzle 7. The element 9 of the nozzle structure with only channel A13 (only with the first channel) is connected to the element 10 of the nozzle structure with both channels placed in it - with channel A13 (with the first channel) and channel B14 (second channel), for example, by solder welding or HIP (hot isostatic pressing) .

As shown in FIG. 3, the connection section 12 is located exclusively in a region in which only one channel A13 (first channel) is formed from both channels, channel A13 (first channel) and channel B14 (second channel). This design prevents the occurrence of thermal stress on the pilot nozzle due to the temperature difference of the fuel or combustion air passing through the corresponding channel, and, in turn, ensures the reliability of section 12.

Here, the aforementioned HIP (hot isostatic pressing) method is preferably used to connect the nozzle element 9 and the nozzle element 10 to each other at the connection portion 12 . The use of the HIP method makes it possible to prevent, to the greatest extent possible, the appearance of non-welded areas and thereby reduce the thermal stress caused by the non-uniformity of the shape in the joint area 12 .

As described above, according to the present invention, it is possible to provide a fuel injector with low thermal stress caused by a temperature difference between the fuel and combustion air that pass through it, and a gas turbine combustor using this fuel injector, as well as the ability to improve reliability. and longevity of the combustion chamber of the gas turbine.

It should be borne in mind that the present invention is not limited to the above embodiments and includes various modifications. For example, the embodiments discussed above have been described in detail in order to explain the present invention in an easy-to-understand manner, and are not necessarily limited to embodiments having all of the structures described. In addition, it is possible to replace a portion of the structure of one embodiment with the structure of another embodiment, and it is also possible to add the structure of one embodiment to the structure of another embodiment. In addition, it is also possible to add/remove/replace some of the designs of each embodiment with other designs.

List of reference positions

1 - compressor;

2 - combustion chamber;

3 - turbine;

4 - insert of the combustion chamber;

5 - transition compartment;

6 - main nozzle;

7 - pilot nozzle;

8 - direction of flow of combustion gases;

9, 10, 11 - nozzle design element;

12 - section of the connection;

13 - channel A;

14 - channel B;

15 - section of the combustion chamber.

Claims (33)

1. A fuel injector containing a plurality of channels includes: the first channel through which the fuel or combustion air passes; and a second channel through which the fuel or combustion air passes and which is different from the first channel, moreover, the fuel injector includes structural elements, and an integral fuel injector structural element of these structural elements constitutes at least the area in which the first channel and the second channel are located. 2. Fuel injector according to claim 1, characterized in that and the first passage and the second passage are divided into a plurality of sections arranged in the circumferential direction of the fuel injector. 3. Fuel injector according to claim 1 or 2, characterized in that combustion air passes through the first channel, and fuel having a lower temperature than the combustion air passes through the second channel. 4. Fuel injector according to claim 1 or 2, characterized in that from the first channel and the second channel, only the first channel is located near the front end of the fuel injector, and the area in which only the first channel is placed is connected to the area in which the first channel and the second channel are placed. 5. Fuel injector according to claim 4, characterized in that the area in which only the first channel is placed is connected to the area in which the first channel and the second channel are placed by welding or hot isostatic pressing (HIP). 6. The combustion chamber of a gas turbine, containing: a combustion chamber liner that substantially constitutes a section of a combustion chamber in which a gaseous mixture of fuel and combustion air is burned; a transition compartment through which combustion gases are directed from the combustion chamber section to the turbine; a pilot nozzle that supplies fuel and combustion air to the combustion chamber section; and a plurality of main injectors that are placed around the pilot injector and supply fuel and combustion air to the combustion chamber section, wherein the pilot nozzle has: the first channel through which the fuel or combustion air passes; and a second passage through which the fuel or combustion air passes and which is different from the first passage, and the pilot injector includes structural elements, and the one-piece pilot injector structural element of these structural elements constitutes at least the area in which the first channel and the second channel are located. 7. The combustion chamber of a gas turbine according to claim 6, characterized in that both the first channel and the second channel are divided into a plurality of channels arranged in the circumferential direction of the pilot nozzle. 8. The combustion chamber of a gas turbine according to claim 6 or 7, characterized in that combustion air passes through the first channel, and fuel having a lower temperature than the combustion air passes through the second channel. 9. The combustion chamber of a gas turbine according to claim 6 or 7, characterized in that of the first channel and the second channel, only the first channel is located near the front end of the pilot nozzle, and the area in which only the first channel is placed is connected to the area in which the first channel and the second channel are placed. 10. The combustion chamber of a gas turbine according to claim 9, characterized in that the area in which only the first channel is placed is connected to the area in which the first channel and the second channel are placed by welding or hot isostatic pressing (HIP).

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