JP2002257344A - Gas turbine combustor - Google Patents

Abstract

PROBLEM TO BE SOLVED: To obtain a gas turbine combustor in which stabilized combustion is ensured even when high temperature air is used through regeneration cycle while lowering NOx. SOLUTION: Downstream side of the opening of a fuel nozzle 28 and an air nozzle 29 in a slow combustion burner 6 is fully filled with diffuse combustion gas because diffuse flame 17 from a pilot burner 5 reaches thereat. When air 15 and fuel 16 is jetted at high speed, a powerful swirl is formed in the diffuse combustion gas and the air 15 and fuel 16 are diluted before being mixed directly. Oxygen density decreases in the air 15 entrained into the high temperature inert gas, i.e., the diffuse combustion gas, and the fuel 16 is decomposed thermally. When they are mixed, a slow combustion region 21 where generation of NOx is retarded through very slow combustion is formed on the downstream side of the opening of the fuel nozzle 28 and the air nozzle 29.

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 a diffusion combustion system and a slow combustion system by heating are combined, and more particularly to a gas turbine combustor in which the inlet air temperature of the combustor exceeds 600.degree. Operate on condition,
In addition, the present invention relates to a gas turbine combustor most suitable for application to a gas turbine that requires reduction of NOx discharged from the gas turbine combustor.

[0002]

2. Description of the Related Art Conventionally, as a typical method for reducing nitrogen oxides (hereinafter referred to as NOx) in exhaust gas from a gas turbine, there is a method in which a diffusion combustion method and a premix combustion method are used in combination. In the case of such a method, first, air and fuel are diffused and mixed using a swirling flow by a pilot burner performing diffusion combustion to form a stable diffusion flame. Next, a diffusion flame is ignited as a pilot flame in a premixed gas generated by mixing air and fuel in the premixer in advance. This makes it possible to stabilize the formation of premixed flames and achieve stable combustion in the entire gas turbine combustor. Also, by increasing the ratio of premixed combustion to diffusion combustion (premixed combustion is By doing so, the proportion of NOx in the exhaust gas can be reduced.

[0003] As an example of a gas turbine combustor (hereinafter simply referred to as a combustor) that performs both the diffusion combustion method and the premix combustion method as described above, there is one shown in FIG. 2 of JP-A-8-86407. .

[0004]

However, the above prior art has the following problems.

[0005] A regeneration cycle system has been proposed for the purpose of improving energy heat efficiency in a gas turbine.
This is equipped with a regenerative heat exchanger so as to pass between the exhaust path of the combustion gas after the turbine is driven and the path of the compressed air from the compressor to the combustor. After the temperature has been increased, it is supplied to the combustor. In this case, the air temperature of the compressed air exceeds about 600 ° C.

[0006] When high-temperature compressed air is supplied by the regeneration cycle method to a combustor that performs the conventional diffusion combustion method and the premix combustion method in combination, the combustion speed of the premixed flame remarkably increases and the premixed flame is increased. The flashback potential into the mixer increases. If the temperature of the compressed air exceeds the self-ignition temperature of the fuel (the self-ignition temperature of liquefied natural gas frequently used as gas turbine fuel is 630 ° C.), self-ignition may occur in the premixer. The occurrence of flashback or spontaneous ignition in these premixers destabilizes combustion and causes NOx
Cause an increase. Therefore, depending on the regeneration cycle, 6
A gas turbine that uses air exceeding 00 ° C and has low NOx
It is considered difficult to use a premixed combustion system for the purpose of realization.

An object of the present invention is to achieve stable combustion even with high-temperature air by a regeneration cycle and to achieve low NO.
An object of the present invention is to provide a gas turbine combustor capable of achieving x conversion.

[0008]

(1) In order to achieve the above object, the present invention provides a pilot burner for burning while diffusing and mixing air and fuel, and air and fuel contained in the combustion gas generated by the pilot burner. A slow-burning burner for injecting the fuel individually, mixing the fuel with the combustion gas, and causing a thermal reaction to perform slow combustion is provided.

By providing the pilot burner and the slow burner in this manner, slow burn can be performed without forming a clear flame and a high-temperature portion using stable high-temperature combustion gas by diffusion combustion. As a result, even when high-temperature compressed air from the regeneration cycle is used, stable combustion is possible without reducing flashback or spontaneous ignition, and NOx can be reduced.

(2) In the gas turbine combustor according to the above (1), preferably, the slow burner is a burner that jets the air and the fuel individually into a jet gas into the combustion gas generated by the pilot burner. There is.

By jetting air and fuel into the combustion gas in the form of a jet, the air and the fuel jetted into the combustion gas mix with the combustion gas and cause a thermal reaction.

(3) In the gas turbine combustor according to the above (1), preferably, the slow burner burns the air and the fuel into the combustion gas generated by the pilot burner at such a high speed that they do not come into direct contact with each other. It shall be a burner that spouts individually.

By injecting the fuel into the combustion gas at such a high speed that the air and fuel do not come into direct contact with each other, the air and fuel injected into the combustion gas mix with the combustion gas and then cause a thermal reaction. Become.

(4) In the gas turbine combustor according to the above (1), preferably, the slow burner burns the air and the fuel at a high speed of at least twice the average outflow speed of the combustion gas at the outlet of the gas turbine combustor. It is assumed that it is a burner spouting.

By ejecting air and fuel at such a speed, the air and fuel ejected into the combustion gas cause a thermal reaction after mixing with the combustion gas.

(5) In the gas turbine combustor according to any one of the above (1) to (4), preferably, the slow burner includes a plurality of nozzle arrays arranged concentrically around the pilot burner. Burner.

(6) In the gas turbine combustor according to any one of the above (1) to (4), preferably, the slow burner includes a plurality of independent slow burners disposed around the pilot burner. Shall have.

(7) In the gas turbine combustor according to any one of the above (1) to (4), preferably, the slow combustion burner has a fuel supply system divided into a plurality of parts, and the fuel is burned. It is assumed that the burner is configured to switch the number of the supplied fuel supply systems stepwise according to the magnitude of the load of the gas turbine.

Thus, the starting and the operation according to the load of the gas turbine can be performed.

[0020]

Embodiments of the present invention will be described below with reference to the drawings.

FIG. 1 is an axial sectional view of a combustor according to a first embodiment of the present invention, and FIG. 2 is a sectional view taken along a line A in FIG.
FIG. 3 is a side view of the combustor as viewed from above. In the following,
The upstream side in the combustion gas flow direction B is defined as the rear side, and the downstream side is defined as the front side.

In FIG. 1, reference numeral 1 denotes a combustor. The combustor 1 has a cylindrical combustor case 2 and a combustor case 2.
And a burner assembly 3 mounted inside one end of the burner. The burner assembly 3 includes a pilot burner 5 located at the center and a slow burning burner 6 surrounding the pilot burner 5. The slow burner 6 has a cylinder 3a and an annular nozzle plate 4 for slow combustion installed at the tip of the cylinder 3a. A pilot burner 5 is installed inside the nozzle plate 4. A partition disk 7 is arranged behind the cylindrical body 3a, and an annular manifold 23 is attached to a rear side surface of the partition disk 7.

A plurality of dilution air holes 8 are formed in the combustor case 2 at a position slightly forward of the center in the axial direction, and are spaced apart in the circumferential direction.

A seal spring 9 is provided on the outer periphery of the burner assembly 3 on the distal end side of the cylindrical body 3a, and the seal spring 9 ensures airtightness of a fitting portion with the combustor case 2.

The slow-burning burner 6 has eight slow-burning nozzles 10 concentrically arranged in the circumferential direction of the nozzle plate 4. Each slow-burning nozzle 10 is located at the center and is located at the front of the nozzle plate 4. A fuel nozzle 28 that opens,
It comprises eight air nozzles 29 formed around the fuel nozzles 28 of the nozzle plate 4. Fuel nozzle 2
Reference numeral 8 denotes an opening at the tip of the fuel nozzle tube 12b.

The pilot burner 5 has one fuel nozzle 11 at the center thereof, and a swirler 13 is coaxially provided on the outer periphery thereof. The fuel nozzle 11 is formed by an opening at the tip of a fuel nozzle tube 12a. The swirler 13 ejects the air that has flowed in from the rear to the front while swirling in the circumferential direction.

The partition disk 7 is arranged at a distance from the rear end of the cylindrical body 3a, and the fuel nozzle pipes 12a and 12b of the pilot burner 5 and each slow burner 6 are connected to the partition disk 7. The fuel nozzle tubes 12a extend further rearward through the partition disk 7, and the fuel nozzle tubes 12b of the slow burner 6 communicate with the inside of the annular manifold 23.

The swirler 13 of the pilot burner 5
Each air nozzle 11 of each slow combustion burner 6 opens to a compressed air chamber 14 which is a space formed between the partition disk 7 and the nozzle plate 4.

FIG. 3 is a schematic configuration diagram of a regeneration cycle type gas turbine using the gas turbine combustor according to the present embodiment.

In FIG. 3, a compressor 101, a turbine section 103, a regenerative heat exchanger 104, and the above-described combustor 1 constitute an entire gas turbine. The combustor 1 burns fuel with the compressed air supplied from the compressor 101 to generate a combustion gas.
6 to the turbine section 103, and the turbine section 103
Drives the generator (generator) G with the combustion gas to convert the energy of the combustion gas into electric energy. The combustion gas that has worked in the turbine section 103 is discharged outside through an exhaust gas passage 107.

The compressor 101 is connected to the compressed air passage 10.
5 passes through the regenerative heat exchanger 104 to recover exhaust heat from exhaust gas passing through the exhaust gas
High-temperature air exceeding 0 ° C. is supplied to the combustor 1.

Returning to FIG. 1, the compressed air 15 supplied from the compressed air passage 105 flows into the swirler 13 of the pilot burner 5 and the air nozzle 29 of each slow burning burner 6 through the compressed air chamber 14 and burns. It is ejected into the container case 2. At this time, the compressed air 15 from the swirler 13
Is ejected while turning in the circumferential direction.

From the start of the gas turbine to the time of low load, the fuel 16 is supplied only from the fuel supply system (not shown) to the fuel nozzle 11 of the pilot burner 5 via the fuel nozzle pipe 12a to form a diffusion flame 17. This diffusion flame 17 is formed stably because the swirling air jet ejected from the swirler 13 burns while directly mixing with the fuel 16. At this time, the compressed air 15 is also ejected from each of the air nozzles 29 of each of the slow combustion nozzles 10 and serves to dilute the diffusion combustion gas.

Thereafter, as the load on the gas turbine increases,
The fuel 1 is also supplied to each fuel nozzle 28 of each slow combustion nozzle 10 via the annular manifold 23 and the fuel nozzle pipe 12b.
6 are supplied. At this time, the fuel 16 is ejected in the form of a high-speed jet as described later, and by using the diffusion flame of the pilot burner 5 as the pilot flame, stable slow combustion is performed and operation at low NOx is maintained. At this time, the inflow of the dilution air 18 from each of the dilution air holes 8 of the combustor case 2 promotes the mixing of the diffusion combustion gas and the slow combustion gas, and adjusts the temperature of the entire combustion gas. Is done.

Next, the principle of slow combustion will be described with reference to FIG. FIG. 4 is an axial cross-sectional view of a main part of the slow combustion burner 6 during slow combustion. The upper part in the figure corresponds to the downstream side of the combustion gas (the front side of the combustor 1).

In FIG. 4, the fuel 16 and the air 15 are individually jetted from the fuel nozzle 28 and the air nozzle 29 in the form of a high-speed jet of about several tens m / s. Once a diffusion flame is formed on the downstream side of the fuel nozzle 28 and the air nozzle 29, the air 15 and the fuel 16 are jetted into the high-temperature combustion gas 20 resulting from the combustion. A strong entrainment vortex 19 of gas 20 is formed. Therefore, the jetted air 15 is diluted by involving and mixing the high-temperature combustion gas 20 which is a kind of inert gas, and the oxygen concentration is relatively reduced.
In addition, the injected fuel 16 entrains and mixes the high-temperature combustion gas 20, so that the heat of the high-temperature combustion gas 20 causes thermal decomposition.

Here, for example, when the oxygen concentration of the air 15 is about 21% of normal and the fuel 16 is methane (CH ₄ ), the following reaction proceeds.・ Air: air (21% O ₂ ) + combustion gas (15% O ₂ ) →
Dilution air (17 to 18% O ₂ ) Fuel: 2CH ₄ + O ₂ → 2CO + 2H ₂ (pyrolysis) The fuel nozzle 28 and the air nozzle 29 are formed by the dilute air having a relatively low oxygen concentration and the pyrolyzed fuel. A thermal reaction region (hereinafter, referred to as a slow combustion region 21) in which a thermal reaction (a kind of combustion) proceeds very slowly as follows:
Is formed.

2H ₂ + O ₂ → 2H ₂ O + reaction heat 2CO + O ₂ → 2CO ₂ + reaction heat In the dead water region between the fuel nozzle 28 and the air nozzle 29, a circulating flow 22 is formed. The stable slow combustion region 21 is formed. Since no definite flame and high-temperature portion are formed in the slow combustion region 21, the generation of NOx can be suppressed.

Here, the air 15 and the fuel 1 in the present invention
The ejection speed of No. 6 is preferably at least twice the average combustion gas flow velocity at the outlet of the gas turbine combustor 1.

If the ejection speed is too slow, the air 15 and the fuel 16
Are directly mixed and then burned immediately afterwards to form a distinct flame or high-temperature portion, thereby increasing the generation of NOx.
Depending on the type of the fuel 16 and other conditions, if the injection speed is at least twice as high as the average combustion gas flow velocity at the gas turbine combustor outlet, the air and fuel are individually injected into the combustion gas without being directly mixed. It has been found that good slow combustion can be performed.

Next, as a comparative example, a conventional combustor using both a diffusion combustion system and a premix combustion system will be described.

FIG. 5 is an axial sectional view of such a conventional combustor. In FIG. 5, reference numeral 200 denotes a pilot burner for diffusion combustion, which includes a diffusion flame fuel nozzle 201 and a swirler 202, and a double-tube premixer 203 is provided around the pilot burner 200. Have been. Compressed air 15 is supplied from the rear of the premixer 203, and a fuel nozzle 204 is provided therein.
Is provided.

In the premixer 203, the compressed air 15 supplied from the rear and the fuel 1 injected from the fuel nozzle 204
6 are mixed with each other to generate a premixed gas 205. This premixed gas 205 is discharged from the front end of the premixer 203,
The diffusion flame 206 by the pilot burner 202 is ignited as a pilot flame to perform premix combustion. This premixed combustion is performed in the premixed gas 20 which is already in a state where it is easy to burn.
5 is always ignited directly, so that a stable premixed flame 207 is formed.

In the case of such a combustion system, it is possible to stabilize the combustion of the entire combustor including the premixed flame 207 and, at the same time, to increase the ratio of the premixed combustion to the diffusion combustion (ie, the premixed combustion). By mainly performing combustion), it is possible to reduce the generation of NOx in the exhaust gas.

However, when the regeneration cycle system is applied to such a combustor for the purpose of improving energy heat efficiency, the compressed air supplied to the combustor exceeds 600 ° C., and the combustion speed of the premixed flame 207 is reduced. Due to the significant rise, the flashback potential into the premixer 203 also increases significantly. Furthermore, if the temperature of the compressed air 15 exceeds the self-ignition temperature of the fuel (the self-ignition temperature of liquefied natural gas used for gas turbine operation is 630 ° C.), autoignition may occur in the premixer 203. When the regeneration cycle method is applied, the premix combustion method cannot be used due to such a problem.

In contrast to such a conventional technique, the above-described combustor 1 according to the present embodiment does not use a premixed gas which is liable to generate spontaneous ignition or flashback, and uses slow combustion which does not form a clear flame and a high-temperature portion. Therefore, even when high-temperature compressed air exceeding 600 ° C. is used, generation of NOx can be reliably reduced without generating flashback or spontaneous ignition.

Therefore, by using the gas turbine combustor according to the present embodiment, it is possible to suppress the generation of NOx while performing stable combustion even when using high-temperature air in the regeneration cycle, and to achieve high energy heat efficiency in the regeneration cycle system. It is possible to operate with.

A second embodiment of the present invention will be described with reference to FIGS. In the figure, parts that are the same as the parts shown in FIGS. 2 and 3 are given the same reference numerals, and descriptions thereof will be omitted. In the present embodiment, a plurality of slow burning nozzles are provided as slow burning burners each having an independent configuration.

6 and 7, the combustor 1A according to the present embodiment includes four slow combustion burners 6A which are fitted to the plate 27 and are formed independently of each other.

Each slow-burning burner 6A has a small-diameter burner case 24 provided with a disk 25 at the end thereof, a disk 25 provided with one fuel nozzle 28 at the center, and eight air nozzles 29 provided around the fuel nozzle 28. One set of slow combustion nozzles 10 is configured.

In this embodiment constructed as described above, diffusion combustion is performed by the pilot burner 5, and all the slow combustion burners 6 are formed via the annular manifold 23.
By supplying the fuel 16 collectively to A, the same slow combustion as in the first embodiment can be performed, and even when the high-temperature compressed air 15 from the regeneration cycle is used, flashback or spontaneous ignition occurs. And the generation of NOx can be reduced.

In this embodiment, since each slow burner 6A is configured independently of the burner assembly 3A (configured separately from the plate 27), it can be individually repaired or replaced. This facilitates maintenance work.

A third embodiment of the present invention will be described with reference to FIGS. In the figure, parts that are the same as the parts shown in FIGS. 2 and 3 are given the same reference numerals, and descriptions thereof will be omitted. In the present embodiment, the number of slow combustion nozzles for supplying fuel according to the magnitude of the turbine load is increased or decreased stepwise with respect to eight sets of slow combustion nozzles to which fuel is supplied from two independent fuel supply systems. It is designed to be switchable.

8 and 9, in a combustor 1B according to the present embodiment, a pilot burner 5 is provided in the center of a nozzle plate 4, and eight sets of slow combustion nozzles 10a, 10b, 10c, 10d, 1
The slow burner 6 is provided by forming 0e, 10f, 10g, and 10h. The fuel nozzle pipes 12b constituting the fuel nozzles 28 of the slow combustion nozzles respectively penetrate the rear partition disk 7 and communicate with the single manifold 26 to be connected to the fuel supply system.

The fuel supply system is divided into two fuel supply systems 31, 32 of a first stage and a second stage.
Four slow combustion nozzles 10a belonging to the stage fuel supply system,
10b, 10c, 10d and the other four slow burning nozzles 10e, 10f, 10g, belonging to the second stage fuel supply system.
10h are alternately arranged in the circumferential direction for each stage fuel supply system. The fuel supply systems 31 and 32 of each stage can independently control the supply amount of the fuel 16 by switching the fuel valves 33 and 34.

Next, an operation method of the combustor configured as described above will be described with reference to FIG. FIG. 10 is a diagram illustrating the relationship between the turbine load and the amount of fuel supplied to each burner. In FIG. 10, when the gas turbine is started (turbine load 0%) and the turbine load is less than X% and the load is low, the fuel 16 is supplied only to the pilot burner 5 and operated only with the stable diffusion flame 17. Then, the fuel supply amount is continuously increased / decreased and adjusted according to the magnitude of the turbine load.

Further, from the time when the gas turbine is started, the compressed air 15 is supplied from the compressed air chamber 14 and injected from the swirler 13 to mix with the fuel 16 and burn.
The compressed air 15 is also ejected from the air nozzle 29 of each slow combustion nozzle 10 to dilute the diffusion combustion gas.

Thereafter, when the turbine load increases and reaches X%, the fuel supply amount to the pilot burner 5 is reduced, and the fuel valve 33 is opened to open the slow combustion nozzles 10a and 10b belonging to the first stage fuel supply system. , 10c, 10d
Supply the required amount of fuel for slow combustion.

Here, as described above, the slow combustion nozzle 1
In order for 0a, 10b, 10c, and 10d to perform slow combustion,
Since the air 15 and the fuel 16 need to be jetted and mixed at such a high speed that they do not burn even if they come into direct contact with the diffusion combustion gas, the diameters of the fuel nozzle 28 and the air nozzle 29 are fixed. Slow combustion can be performed only when a large amount of fuel is required.

That is, the turbine load of this X% corresponds to four slow combustion nozzles 1 belonging to the first stage fuel supply system 31.
The partial load requiring the minimum output (the output at the time of the minimum fuel supply at which the slow combustion can be performed) in the case where the turbines 0a, 10b, 10c, and 10d are operated together is also a large load. For the first time, the slow combustion nozzles 10a, 10b belonging to the first stage fuel supply system 31
The same amount of fuel 16 is supplied to 10c and 10d, and slow combustion is performed in half of the entire slow burner 6.

When the turbine load is a medium load between X% and Y%, the fuel supply amount to the first stage fuel supply system 31 is kept constant, and the fuel is supplied to the pilot burner 5 according to the magnitude of the turbine load. Is adjusted by continuously increasing or decreasing the fuel supply amount.

When the turbine load further increases and reaches Y%, the fuel supply amount to the pilot burner 5 is reduced again, and the fuel valve 34 is opened to open the slow combustion nozzle 10e belonging to the second stage fuel supply system 32. , 10f, 10
g, start supplying the required amount of fuel to 10h,
Slow combustion is performed by the entire slow burner 6.

When the turbine load is at a high load between Y% and 100% rated load, the fuel supply to the pilot burner 5 is kept constant at a low level. Second stage fuel supply system 31,3
2 all slow combustion nozzles 10a, 10b, 10
The fuel supply amounts to c, 10d, 10e, 10f, 10g, and 10h are continuously increased or decreased to be adjusted.

According to the present embodiment and the fuel supply method configured as described above, the fuel supply amount is increased and decreased appropriately in response to any partial load between the start of the gas turbine and the rated load of the turbine. However, each slow burning nozzle 10
a, 10b, 10c, 10d, 10e, 10f, 10
It is possible to always maintain good slow combustion by appropriately maintaining the fuel ejection speed at g and 10h.

In the third embodiment, eight slow combustion nozzles 10a, 10b, 10c, 10d,
Half (four) of each of 10e, 10f, 10g, and 10h is distributed to two fuel supply systems 31, 32, and the fuel 1
Although the number of fuel supply systems for supplying 6 is increased or decreased in two stages, the gas turbine combustor of the present invention is not limited to this. For example, as shown in FIGS. 11 and 12, four slow combustion burners 6Aa, 6Ab, 6Ac, 6Ad and four independent fuel supply systems 35, 3 for each slow burner.
6, 37, and 38, each of which gradually
It is also possible to adopt a configuration in which supply is performed. In this case, the relationship between the turbine load and the amount of fuel supplied to each burner is such that fuel is supplied in four stages for each fuel supply system as shown in FIG. With this configuration, it is possible to supply an amount of fuel appropriately corresponding to the change in the partial load as compared with the configuration in which the two-stage fuel supply is performed, and to operate while maintaining better slow combustion. Costs can be reduced.

[0066]

According to the present invention, it is possible to perform slow combustion without forming a clear flame and a high-temperature portion by using stable high-temperature combustion gas by diffusion combustion. Even when high-temperature compressed air is used, stable combustion is possible without reducing flashback or spontaneous ignition, and NOx can be reduced.

Further, according to the present invention, while appropriately increasing or decreasing the fuel supply amount in response to any partial load, the fuel injection speed in each slow combustion burner is appropriately maintained, and a good slow combustion is always achieved. Can be realized.

[Brief description of the drawings]

FIG. 1 is a schematic configuration diagram of an example of a regeneration cycle type gas turbine to which a combustor according to the present invention is applied.

FIG. 2 is an axial sectional view of the combustor according to the first embodiment of the present invention.

FIG. 3 is a side view of the combustor as viewed from an arrow A in FIG. 2;

FIG. 4 is an axial cross-sectional view of a slow burning burner during slow burning;

FIG. 5 is an axial sectional view of a combustor using both a diffusion combustion system and a premix combustion system.

FIG. 6 is an axial sectional view of a combustor according to a second embodiment of the present invention.

FIG. 7 is a side view of the combustor as viewed from an arrow A in FIG.

FIG. 8 is an axial sectional view of a combustor according to a third embodiment of the present invention.

FIG. 9 is a side view of the combustor as viewed from an arrow A in FIG.

FIG. 10 is a diagram showing a relationship between a turbine load of a combustor and a fuel supply amount to each burner according to a third embodiment of the present invention.

FIG. 11 is an axial cross-sectional view of a combustor according to a modification of the third embodiment of the present invention.

FIG. 12 is a side view of the combustor as viewed from an arrow A in FIG. 11;

FIG. 13 is a diagram illustrating a relationship between a turbine load of a combustor and a fuel supply amount to each burner according to a modification of the third embodiment of the present invention.

[Explanation of symbols]

 1, 1A, 1B, 1C Combustor 2 Combustor case 3, 3A Burner assembly 3a Tube 4 Nozzle plate 5 Pilot burner 6, 6A, Slow combustion burner 7, 7A Partition disk 8 Dilution air hole 9 Seal spring 10 Slow Combustion nozzle 11 Fuel nozzle 12a, 12b Fuel nozzle tube 13 Swirler 14 Compressed air chamber 15 Compressed air 16 Fuel 17 Diffusion flame 18 Dilution air 19 Entrained vortex 20 High temperature combustion gas 21 Slow combustion area 22 Circulating flow 23 Annular manifold 24 Burner case 25 disk 26 unitary manifold 27 plate 28 fuel nozzle 29 air nozzle 31 first stage fuel supply system 32 second stage fuel supply system 33,34 fuel valve 35 first stage fuel supply system 36 second stage fuel supply system 37 third Stage fuel supply system 38 4th stage fuel supply Integrated 39, 40, 41, 42 fuel valve 101 compressor 102 combustor 103 turbine unit 104 regenerative heat exchanger 105 the compressed air passage 106 combustion gas passage 107 exhaust passage G Generator

 ──────────────────────────────────────────────────続 き Continuing on the front page (72) Inventor Hiromi Koizumi 7-2-1, Omika-cho, Hitachi City, Ibaraki Prefecture Inside Power & Electric Development Laboratory, Hitachi, Ltd. (72) Inventor Narika Kobayashi Omika, Hitachi City, Ibaraki Prefecture 7-2-1, Machi-cho F-term in Hitachi, Ltd. Electric Power and Electrical Development Laboratory 3K019 AA06 BA04 BB02 CC01 3K065 TA01 TA14 TC08 TD05 TG04 TH01 TH19 TJ03 TJ06 TL03

Claims (7)

[The claims] 1. A pilot burner that burns while diffusing and mixing air and fuel, and air and fuel are separately jetted into combustion gas generated by the pilot burner, and mixed with the combustion gas, respectively, to perform a thermal reaction. A gas turbine combustor comprising: a slow combustion burner for performing slow combustion by causing the burner to generate a slow combustion. 2. The gas turbine combustor according to claim 1, wherein the slow burner is a burner that jets the air and the fuel individually into a jet gas into the combustion gas generated by the pilot burner. And gas turbine combustor. 3. The gas turbine combustor according to claim 1, wherein said slow burner individually blows said air and fuel into the combustion gas generated by said pilot burner at such a high speed that they do not come into direct contact. A gas turbine combustor characterized by being a burner. 4. The gas turbine combustor according to claim 1, wherein said slow combustion burner injects said air and fuel at a high speed of at least twice the average outflow velocity of the combustion gas at an outlet of said gas turbine combustor. A gas turbine combustor characterized by the following. 5. The gas turbine combustor according to claim 1, wherein said slow burner is a burner having a plurality of nozzle arrangements arranged concentrically around said pilot burner. A gas turbine combustor characterized in that: 6. The gas turbine combustor according to claim 1, wherein said slow burner has a plurality of independent slow burners disposed around said pilot burner. And gas turbine combustor. 7. The gas turbine combustor according to claim 1, wherein said slow combustion burner has a fuel supply system divided into a plurality of fuel supply systems, and said fuel supply system supplies said fuel. A gas turbine combustor, characterized in that the number of burners is switched stepwise according to the load of the gas turbine.