JP2010249449A - Pilot combustion burner for gas turbine - Google Patents

Abstract

<P>PROBLEM TO BE SOLVED: To drastically reduce vibration caused by flowing of compressed air and prevent blow off at start. <P>SOLUTION: This pilot combustion burner for a gas turbine includes: a pilot fuel nozzle 31 inside which a plurality of fuel flow passages 51 for premixed combustion and a plurality of fuel flow passages 52 for diffusing combustion are formed independently; a pilot burner cylinder 32 concentric with respect to the pilot combustion nozzle 31 and having an end on the upstream side arranged to surround an end on the downstream side of the pilot combustion nozzle 31; and a plurality of turning blades 33 radially arranged at the end on the downstream side of the pilot combustion nozzle 31 and applying turning force to compressed air passing through a ring-shaped air passage 34 formed between the end on the downstream side of the pilot combustion nozzle 31 and the end on the upstream side of the pilot burner cylinder 32 to make the compressed air form a turning air flow. <P>COPYRIGHT: (C)2011,JPO&INPIT

Description

  The present invention relates to a pilot combustion burner for a gas turbine arranged at the center of a combustor of a gas turbine.

As a pilot combustion burner of a gas turbine arranged at the axial center of a combustor of a gas turbine, for example, the one disclosed in Patent Document 1 is known.
However, the pilot combustion burner of the gas turbine disclosed in Patent Document 1 is configured to perform only diffusion combustion in which a local high temperature portion is generated during combustion, and there is a limit to reducing NOx. is there.
Therefore, in recent years, a pilot combustion burner for a gas turbine configured to perform premixed combustion in which a local high temperature portion does not occur during combustion has been proposed (see, for example, Patent Document 2).

JP 2004-101081 A US Pat. No. 6,848,260

  However, the pilot combustion burner of the gas turbine disclosed in Patent Document 2 is provided with a ring-shaped manifold (cavity) 74 on the outer peripheral side (radially outer side) of the swirler vane (swivel blade) 78. The fuel 96 supplied to is injected through an outlet (injection hole) 84. For this reason, a manifold 74 as an additional material is provided in the passage of compressed air for combustion, which may cause fluid resistance. That is, the manifold 74 is vibrated by the compressed air toward the swirler vane 78, and not only does the breakage due to vibration occur in the connection portion between the fuel pipe 72 and the manifold 74 or the connection portion between the manifold 74 and the swirler vane 78, but also the air flow. Turbulence and pressure loss may affect stable combustion.

  The present invention has been made in view of the above circumstances, and can greatly reduce vibrations caused by the flow of compressed air, and can prevent blow-off during startup. An object is to provide a pilot combustion burner for a gas turbine.

The present invention employs the following means in order to solve the above problems.
A pilot combustion burner of a gas turbine according to the present invention is a pilot combustion burner of a gas turbine arranged at the axial center of a combustor of a gas turbine, and is used for a plurality of premixed combustion in the inside along an axial direction. A pilot combustion nozzle in which a fuel flow path and a plurality of diffusion combustion fuel flow paths are independently formed, concentric with the pilot combustion nozzle, and an end on the upstream side of the pilot combustion nozzle A pilot burner cylinder disposed in a state of surrounding a downstream end of the pilot combustion nozzle; and a downstream end of the pilot combustion nozzle disposed radially at the downstream end of the pilot combustion nozzle; , A plurality of swirl blades for applying a swirl force to the compressed air passing through a ring-shaped air passage formed between the upstream end portion of the burner cylinder and turning the compressed air into a swirl air flow It is equipped with a.

According to the pilot combustion burner of the gas turbine according to the present invention, the fuel flow path for premixed combustion is drilled in the pilot combustion nozzle, and swirls from the inner peripheral side (radially inner side: blade root side) of the swirl blade. Since the fuel is supplied to the blade, compared with the one disclosed in Patent Document 2 described above, the portion that vibrates due to the flow of compressed air is greatly reduced, and the compressed air flows. The vibration caused by can be greatly reduced.
Also, at startup when combustion is unstable, diffuse combustion is performed using only the fuel passage for diffusion combustion, and when the gas turbine reaches a predetermined load, only the fuel passage for premix combustion is used. By performing the premixed combustion, it is possible to prevent blow-off at the time of start-up and to achieve low NOx.

  In the pilot combustion burner of the gas turbine, it is more preferable that the fuel flow path for premixed combustion is formed of an upstream flow path and a downstream flow path having a smaller sectional area.

In the pilot combustion burner of the gas turbine, it is more preferable that a cavity that is continuous in the circumferential direction is provided in the middle of the fuel flow path for premixed combustion.
Furthermore, it is more preferable that the cavity is provided at a transition portion between the upstream flow path and the downstream flow path. Or the said cavity may be provided in the blade root part of the said swirl | wing blade.

According to such a pilot combustion burner of a gas turbine, the number of fuel channels for premixed combustion to the cavity can be minimized, and the inner diameter of each fuel channel from the cavity to the swirl vane can be reduced. The outer diameter of the downstream end portion of the pilot combustion nozzle can be reduced.
This also makes it possible to minimize the pressure loss of the fuel passing through the fuel flow path.

  In the pilot combustion burner of the gas turbine, the radially outer side of the downstream end of the pilot combustion nozzle located downstream from the blade root of the swirl blade is covered, and the air passage is connected to the air on the inner peripheral side. It is more preferable that a partition wall that partitions the passage and the outer air passage is provided.

  According to such a pilot combustion burner of a gas turbine, the end portion on the downstream side of the pilot combustion nozzle is covered with a cylindrical air layer (film air) that has passed through the air passage on the inner peripheral side during premix combustion. Therefore, damage due to high temperature at the downstream end of the pilot combustion nozzle can be prevented.

  A combustor for a gas turbine according to the present invention includes a pilot combustion burner for a gas turbine that can significantly reduce vibrations caused by the flow of compressed air and that can prevent blowout during startup. is doing.

According to the combustor of the gas turbine according to the present invention, it is possible to prevent the operation of the gas turbine from being stopped due to breakage or damage of component parts and blowout, and to improve the reliability.
Also, at startup when combustion is unstable, diffuse combustion is performed using only the fuel passage for diffusion combustion, and when the gas turbine reaches a predetermined load, only the fuel passage for premix combustion is used. Thus, NOx can be reduced by performing premixed combustion.

  The gas turbine according to the present invention can prevent the gas turbine from shutting down due to breakage or damage of components or blown off, and can improve the reliability of the gas turbine. NOx is produced by performing diffusion combustion using only the fuel flow path for combustion and, when the gas turbine reaches a predetermined load, performing premix combustion using only the fuel flow path for premix combustion. The gas turbine combustor can be reduced.

  According to the gas turbine of the present invention, the reliability can be improved and the reduction of NOx can be realized.

  According to the pilot combustion burner of the gas turbine according to the present invention, it is possible to significantly reduce the vibration caused by the flow of the compressed air and to prevent the blow-off at the start-up.

It is a schematic block diagram which shows the combustor of the gas turbine provided with the pilot combustion burner of the gas turbine which concerns on 1st Embodiment of this invention. It is a figure which shows the combustor of the gas turbine shown in FIG. 1, Comprising: It is a perspective view which decomposes | disassembles and shows a fuel nozzle, an inner cylinder, and a tail cylinder. It is sectional drawing which expanded and showed the principal part of FIG. FIG. 4 is a cross-sectional view taken along arrow IV-IV in FIG. FIG. 5 is a plan view of a swirling blade viewed along an arrow A shown in FIG. 4. It is VI-VI arrow sectional drawing of FIG. It is a VII-VII arrow sectional view of Drawing 3. It is sectional drawing which expanded and showed the principal part of the pilot combustion burner of the gas turbine which concerns on 2nd Embodiment of this invention. It is IX-IX arrow sectional drawing of FIG.

Hereinafter, the pilot combustion burner of the gas turbine which concerns on 1st Embodiment of this invention is demonstrated, referring FIGS. 1-7.
FIG. 1 is a schematic configuration diagram showing a combustor of a gas turbine equipped with a pilot combustion burner of a gas turbine according to the present embodiment, and FIG. 2 is a diagram showing the combustor of the gas turbine shown in FIG. 3 is an exploded perspective view showing the inner cylinder and the tail cylinder, FIG. 3 is an enlarged cross-sectional view showing the main part of FIG. 1, FIG. 4 is a cross-sectional view taken along the line IV-IV in FIG. 3, and FIG. FIG. 6 is a cross-sectional view taken along arrow VI-VI in FIG. 3, and FIG. 7 is a cross-sectional view taken along arrow VII-VII in FIG.

  A gas turbine (not shown) including a combustor (hereinafter referred to as “combustor”) 10 of the gas turbine shown in FIGS. 1 and 2 includes a compressor (not shown) in addition to the combustor 10. A turbine (not shown). Many gas turbines have a plurality of combustors 10, and the air (compressed air) compressed by the compressors and the fuel supplied to the combustors 10 are mixed and burned in each combustor 10. To generate high-temperature combustion gas. This high-temperature combustion gas is supplied to the turbine to rotate the turbine.

  As shown in FIG. 1, a plurality of combustors 10 are annularly arranged in a combustor casing 11 (only one is shown in FIG. 1). The combustor casing 11 and the gas turbine casing 12 are filled with compressed air to form a passenger compartment 13. Air that has been compressed by a compressor is introduced into the passenger compartment 13. The introduced compressed air enters the inside of the combustor 10 from an air inlet 14 provided in the upstream portion of the combustor 10. Inside the inner cylinder 15 of the combustor 10, the fuel supplied from the combustion burner 16 and the compressed air are mixed and burned. The combustion gas generated by the combustion is supplied to the turbine chamber side through the tail cylinder 17 and rotates a turbine rotor (not shown).

FIG. 2 is a perspective view showing the combustion burner 16, the inner cylinder 15, and the tail cylinder 17 separately.
As shown in FIG. 2, the combustion burner 16 includes a plurality of main combustion burners 18 of a gas turbine and one pilot combustion burner 19.
The plurality of main combustion burners 18 are arranged inside the inner cylinder 15 and so as to surround the periphery of the pilot combustion burner 19 as shown in FIG. The fuel injected from the main combustion burner 18 is premixed with the swirled air by the swirler vanes (swirler vanes) 20 (see FIG. 2) of the main combustion burner 18 and burns inside the inner cylinder 15.

The main combustion burner 18 includes a main fuel nozzle (hereinafter referred to as “main nozzle”) 21, a main burner cylinder 22, and swirl vanes 20 as main elements.
The main burner cylinder 22 is concentric with the main nozzle 21 and is disposed so as to surround the main nozzle 21. For this reason, a ring-shaped air passage (not shown) is formed between the outer peripheral surface of the main nozzle 21 and the inner peripheral surface of the main burner cylinder 22, and this air passage extends from the upstream side to the downstream side. Opposite, compressed air (not shown) flows.

A plurality of swirling blades 20 (six in this embodiment) are arranged radially from the outer peripheral surface of the main nozzle 21 and along the axial direction of the main nozzle 21.
Each swirl vane 20 imparts a swirl force to the compressed air flowing through the air passage formed between the outer peripheral surface of the main nozzle 21 and the inner peripheral surface of the main burner cylinder 22, and this swirl air flow It is to make.

  A plurality of (for example, two) injection holes (not shown) penetrating in the thickness direction are formed on the blade back surface (not shown) of each swirl blade 20, and the blade vent surface (see FIG. A plurality of (for example, two) injection holes (not shown) penetrating in the thickness direction are formed in the not shown. In addition, fuel is supplied to each injection hole via a fuel passage (not shown) formed inside the swirl vane 20 and a fuel passage (not shown) formed inside the main nozzle 21. It has become. The fuel injected from each injection hole is mixed with compressed air to become fuel gas, which is sent to the internal space of the inner cylinder 15 and combusted.

As shown in FIG. 1 or FIG. 3, the pilot combustion burner 19 includes a pilot combustion nozzle (hereinafter referred to as “pilot nozzle”) 31, a pilot burner cylinder 32 (see FIG. 2), swirl vanes (swirler vanes). 33 (refer to FIG. 3).
The pilot burner cylinder 32 is concentric with the pilot nozzle 31, and its base end (left end in FIG. 3) surrounds the tip end (right end in FIG. 3) of the pilot nozzle 31. It is arranged in the state to do. For this reason, a ring-shaped air passage 34 is formed between the outer peripheral surface 31a of the pilot nozzle 31 and the inner peripheral surface 32a of the base end portion of the pilot burner cylinder 32, and this air passage 34 has an upstream side ( Compressed air (not shown) flows from the left side in FIG. 3 to the downstream side (right side in FIG. 3).
In order to simplify the drawing, the swirl vane 33 is not shown in FIG.

A plurality of swirl vanes 33 (eight in this embodiment) are arranged radially from the outer peripheral surface 31 a of the pilot nozzle 31 and along the axial direction of the pilot nozzle 31. Each swirl vane 33 is connected to the pilot nozzle 31 via the blade root portion 35 standing on the outer peripheral surface 31a of the pilot nozzle 31 (extending radially outward from the outer peripheral surface 31a of the tip portion) and the blade root portion 35. And a blade body portion 36 to be connected.
As shown in FIG. 5, each swirl vane 33 is a streamlined member having a planar airfoil shape, and includes a distal end outer peripheral surface 31 a of the pilot nozzle 31 and a proximal end inner peripheral surface 32 a of the pilot burner cylinder 32. A swirling force is applied to the compressed air flowing through the air passage 34 formed therebetween, and this compressed air is turned into a swirling air flow.

  As shown in FIG. 3, a gap 37 is provided between the end surface 36 a on the inner peripheral side (radially inner side: the side close to the pilot nozzle 31) of each blade main body portion 36 and the tip end outer peripheral surface 31 a of the pilot nozzle 31. In the gap 37, one end of a partition wall 38 that partitions the air passage 34 into an inner air passage 34a and an outer air passage 34b is inserted. The partition wall 38 is a hollow cylindrical member extending along the axial direction of the pilot nozzle 31, and a plurality of (for example, four) ribs extending radially (outward in the radial direction) from the outer peripheral surface 31 a of the pilot nozzle 31. 39 is connected (coupled) to the outer peripheral surface 31 a of the tip of the pilot nozzle 31. The other end of the partition wall 38 extends to the tip end surface 31 b of the pilot nozzle 31. That is, the tip end portion of the pilot nozzle 31 located downstream of the rear edge of the blade root portion 35 is covered (enclosed) with the partition wall 38 on the radially outer side.

  A plurality of (for example, two) injection holes (not shown) penetrating in the thickness direction are formed in the blade back surface 33a of each swirl vane 33, and the blade belly surface 33b of each swirl vane 33 has a wall thickness. A plurality of (for example, two) injection holes (not shown) penetrating in the direction are formed. In addition, fuel is supplied to each injection hole via a fuel passage 40 formed inside the swirl vane 33. The fuel injected from each injection hole is mixed with compressed air to become fuel gas, which is sent to the inner space of the inner cylinder 15 through the air passage 34b on the outer peripheral side and combusted.

The pilot nozzle 31 includes two independent fuel flow paths, that is, a fuel flow path 51 for premixed combustion and a fuel flow path 52 for diffusion combustion.
The premixed combustion fuel flow path (hereinafter referred to as “first fuel flow path”) 51 extends in the axial direction to an upstream fuel flow extending to a cavity 53 formed in the vicinity of the upstream side of the blade root portion 35. The passage 54 communicates with the fuel passage 40 having an inner diameter smaller than the inner diameter of the upstream fuel passage 54 and formed inside the cavity 53 and the swirl vane 33 (more specifically, inside the blade root portion 35). And a downstream fuel passage 55.

As shown in FIG. 6, a plurality (four in this embodiment) of the upstream fuel flow paths 54 are provided at equal intervals (90 ° in this embodiment) along the circumferential direction.
As shown in FIG. 7, the cavity 53 is disposed at a transition portion between the upstream fuel flow path 54 and the downstream fuel flow path 55, is continuous along the circumferential direction, and includes a plurality of upstream fuel flow paths 54. It is a ring-shaped space communicating with one end (or one end).
Also, as shown in FIG. 4, a plurality of downstream fuel passages 55 (eight in this embodiment) are provided at equal intervals (45 degrees in this embodiment) along the circumferential direction. One end (downstream end) communicates with a fuel passage 40 formed inside the swirl vane 33 (more specifically, inside the blade root portion 35).

  As shown in FIG. 3, the diffusion combustion fuel flow path (hereinafter referred to as “second fuel flow path”) 52 extends in the axial direction to the upstream vicinity of the cavity 53 to the upstream side fuel flow path. 56 and an inner diameter smaller than the inner diameter of the upstream fuel flow path 56, one end portion (or one end) thereof communicates with the upstream fuel flow path 56, and the other end is opened to the front end surface 31 b of the pilot nozzle 31. And a downstream fuel flow path 57.

As shown in FIG. 6, the upstream fuel flow path 56 is equidistant from the adjacent first fuel flow path 51 (45 degrees in the present embodiment) and is equally spaced along the circumferential direction (this implementation). In the embodiment, a plurality of (four in this embodiment) are provided at intervals of 90 degrees. Moreover, as shown in FIG. 6, these upstream fuel flow paths 54 and 56 are alternately arranged on the same circumference.
As shown in FIG. 7, the downstream fuel flow path 57 is formed on the inner peripheral side (radially inner side) of the cavity 53.
Further, as shown in FIG. 4, the downstream fuel passage 57 is formed on the inner peripheral side (radially inward) of the upstream fuel passage 55 of the first fuel passage 51.
The fuel that has passed through the second fuel flow path 52 is ejected from the opening 58 of the downstream fuel flow path 57 formed in the tip end surface 31 b of the pilot nozzle 31 and mixed with the compressed air that has passed through the air passage 34. As a result, fuel gas is sent to the inner space of the inner cylinder 15 and burned.

Each fuel flow path 54, 55, 56, 57 may be formed by machining such as drilling, or may be formed as a casting hole simultaneously with the main body of the pilot nozzle 31.
Further, in this embodiment, the upstream fuel flow paths 54 and 56 have the same inner diameter and the same number, but the inner diameter and the number may be changed depending on the required fuel flow rate. You may change the distance from. Moreover, each fuel flow path 54,55,56,57 does not need to be a circular cross section, for example, may have a rectangular cross section and a fan-shaped cross section, and is not limited to this embodiment.

According to the pilot combustion burner 19 of the gas turbine according to the present embodiment, the fuel flow path 51 for premixed combustion is formed in the pilot combustion nozzle 31 and the inner peripheral side (radially inner side: blade in the radial direction) of the swirl blade 33. Since the fuel is supplied to the swirl blade 33 from the root portion 35 side), the portion that vibrates due to the flow of compressed air is greatly reduced compared to that disclosed in Patent Document 2 described above. The vibration caused by the flow of compressed air can be greatly reduced.
Further, at the time of start-up where combustion is unstable, only the diffusion combustion fuel flow path 52 is used to perform diffusion combustion. When the gas turbine reaches a predetermined load, only the premixed combustion fuel flow path 51 is used. By using it and performing premix combustion, it is possible to prevent blow-off at the time of start-up and to achieve low NOx.

Further, according to the pilot combustion burner 19 of the gas turbine according to the present embodiment, the cavity 53 that is continuous in the circumferential direction is provided in the middle of the fuel flow path 51 for premixed combustion. The number of mixed combustion fuel flow paths 51, that is, the number of upstream fuel flow paths 54 can be minimized, and the inner diameter of each downstream fuel flow path 55 from the cavity 53 to the swirl vane 33 can be reduced. And the outer diameter of the tip of the pilot combustion nozzle 31 can be reduced.
Furthermore, this makes it possible to minimize the pressure loss of the fuel passing through the fuel flow path 51.

  Furthermore, according to the pilot combustion burner 19 of the gas turbine according to the present embodiment, the radially outer side of the tip end portion of the pilot combustion nozzle 31 located downstream of the blade root portion 35 of the swirl blade 33 is covered, and the air A partition wall 38 is provided to partition the passage 34 into an inner peripheral air passage 34a and an outer peripheral air passage 34b, and a cylindrical air layer that has passed through the inner peripheral air passage 34a during premix combustion. Since the tip portion of the pilot combustion nozzle 31 is covered with (film air), damage to the tip portion of the pilot combustion nozzle 31 due to high temperature can be prevented.

According to the gas turbine combustor 10 provided with the pilot combustion burner 19 of the gas turbine according to the present embodiment, it is possible to prevent the operation of the gas turbine from being stopped due to breakage or damage of components or blow-off, and its reliability. Can be improved.
Further, at the time of start-up where combustion is unstable, only the diffusion combustion fuel flow path 52 is used to perform diffusion combustion. When the gas turbine reaches a predetermined load, only the premixed combustion fuel flow path 51 is used. By using it and performing premix combustion, NOx can be reduced.

  According to the gas turbine provided with the pilot combustion burner 19 of the gas turbine according to the present embodiment, it is possible to improve its reliability and to achieve low NOx.

A pilot combustion burner for a gas turbine according to a second embodiment of the present invention will be described with reference to FIGS.
FIG. 8 is an enlarged cross-sectional view of the main part of the pilot combustion burner of the gas turbine according to the present embodiment, and FIG. 9 is a cross-sectional view taken along arrow IX-IX in FIG.
In order to simplify the drawing, the swirl vane 20 is not shown in FIG.

  As shown in FIG. 8, the pilot combustion burner 61 according to the present embodiment is provided with a cavity 62 instead of the cavity 53, the upstream fuel flow path 54 of the first fuel flow path 51, and the first fuel flow. The downstream fuel flow path 55 of the path 51 is different from that of the first embodiment described above in that it is directly connected (communicated) without passing through the cavity 53. Since other components are the same as those of the first embodiment described above, description of these components is omitted here.

As shown in FIGS. 8 and 9, the blade root portion 35 of the swirl vane 33 according to the present embodiment is formed continuously along the circumferential direction, and the blade body portion 36 of the swirl vane 33 according to the present embodiment. Is erected on the outer peripheral surface 35a of the blade root portion 35 (extends radially outward from the outer peripheral surface 35a of the blade root portion 35).
A cavity 62 that is continuous along the circumferential direction is provided on the outer peripheral side of the blade root portion 35, and this cavity 62 and one end of the downstream fuel flow path 55 are connected via a fuel flow path 63 that extends in the radial direction. Connected (communication). The cavity 62 is connected (communication) to the fuel flow path 40 formed inside the blade main body 36, and the fuel supplied to the fuel flow path 40 is drilled in the blade back surface 33a and the blade belly surface 33b. The fuel injected through each injection hole 64 and injected from each injection hole 64 is mixed with compressed air to become fuel gas, and is sent to the inner space of the inner cylinder 15 through the outer air passage 34b. Will burn.

In this embodiment, a partition wall 65 is provided instead of the partition wall 38 and the rib 39, and an air hole 66 is provided in the blade root portion 35.
The partition wall 65 is a hollow cylindrical member extending along the axial direction of the pilot nozzle 31, one end of which is connected (connected) to the rear edge of the blade root portion 35, and the other end is the tip surface of the pilot nozzle 31. It extends to 31b. The outer peripheral surface 65 a of the partition wall 65 is formed so as to be flush with the outer peripheral surface 35 a of the blade root portion 35, and the inner peripheral surface 65 b of the partition wall 65 is formed so as to be located outside the air hole 66. Yes. That is, the distal end portion of the pilot nozzle 31 located downstream of the rear edge of the blade root portion 35 is covered (enclosed) with the partition wall 65 on the radially outer side.

The air hole 66 is a through hole that penetrates the blade root portion 35 in the length direction (left-right direction in FIG. 8), and on the outer peripheral side (radially outer side) of the second fuel channel 52 as shown in FIG. A plurality (eight in the present embodiment) are provided on the inner circumferential side (radially inward) of the cavity 62 at equal intervals (45 degrees in the present embodiment) along the circumferential direction.
The compressed air or combustion gas that has passed through the blade body 36 is sent to the inner space of the inner cylinder 15 through the outer air passage 34b, and the compressed air that has passed through the air hole 66 is the inner peripheral air. It will be sent to the internal space of the inner cylinder 15 through the passage 34a.
As with the fuel flow paths 54, 55, 56, and 57, the air hole 66 may not have a circular cross section, and may have, for example, an elliptical cross section, a rectangular cross section, a sector cross section, or the like, and is limited to this embodiment. It is not done.

  The effect of the pilot combustion burner 61 according to the present embodiment, the effect of the combustor 10 of the gas turbine including the pilot combustion burner 61 according to the present embodiment, and the gas turbine including the pilot combustion burner 61 according to the present embodiment Since the operational effects are the same as those described in the first embodiment, the description thereof is omitted here.

  The present invention is not limited to the above-described embodiments, and various changes and modifications can be made without departing from the gist of the present invention.

DESCRIPTION OF SYMBOLS 10 Combustor 19 Pilot combustion burner 31 Pilot combustion nozzle 32 Pilot burner cylinder 33 Swivel blade 34 Air passage 34a Inner air passage 34b Outer air passage 35 Blade root portion 38 Partition wall 51 Fuel passage 52 for premix combustion Fuel flow path 53 for diffusion combustion Cavity 61 Pilot combustion burner 62 Cavity 65 Partition wall

Claims (9)

A gas turbine pilot combustion burner arranged at the center of a gas turbine combustor,
A pilot combustion nozzle in which a plurality of premixed combustion fuel flow paths and a plurality of diffusion combustion fuel flow paths are formed independently along the axial direction;
A pilot burner cylinder that is concentric with the pilot combustion nozzle and is disposed in a state in which an upstream end thereof surrounds an end on the downstream side of the pilot combustion nozzle;
A ring-shaped air passage disposed radially at the downstream end of the pilot combustion nozzle and formed between the downstream end of the pilot combustion nozzle and the upstream end of the burner cylinder A pilot combustion burner for a gas turbine, comprising: a plurality of swirling blades that apply a swirling force to the compressed air passing through the swirl to turn the compressed air into a swirling air flow.   The pilot combustion burner for a gas turbine according to claim 1, wherein the fuel flow path for premixed combustion is formed of an upstream flow path and a downstream flow path having a smaller cross-sectional area. .   The pilot combustion burner for a gas turbine according to claim 1 or 2, wherein a cavity continuous in the circumferential direction is provided in the middle of the fuel channel for premixed combustion.   The pilot combustion burner for a gas turbine according to claim 3, wherein the cavity is provided at a transition portion between the upstream flow path and the downstream flow path.   The pilot combustion burner for a gas turbine according to claim 3, wherein the cavity is provided in a blade root portion of the swirl blade.   Covers the radially outer side of the downstream end of the pilot combustion nozzle located downstream from the blade root of the swirl blade, and the air passage includes an inner peripheral air passage and an outer peripheral air passage. A pilot combustion burner for a gas turbine according to any one of claims 1 to 5, wherein a partition wall for partitioning is provided.   A combustor for a gas turbine, comprising the pilot combustion burner for a gas turbine according to any one of claims 1 to 6.   A gas turbine comprising the combustor of the gas turbine according to claim 7. A pilot combustion nozzle in which a plurality of premixed combustion fuel flow paths and a plurality of diffusion combustion fuel flow paths are independently formed along the axial direction,
A pilot burner cylinder that is concentric with the pilot combustion nozzle and is disposed in a state in which an upstream end thereof surrounds an end on the downstream side of the pilot combustion nozzle;
A ring-shaped air passage disposed radially at the downstream end of the pilot combustion nozzle and formed between the downstream end of the pilot combustion nozzle and the upstream end of the burner cylinder And a plurality of swirl vanes that impart a swirling force to the compressed air passing through the swirl to turn the compressed air into a swirling air flow, and are arranged at the axis of the combustor of the gas turbine. A method of operating a combustion burner,
At startup, diffusion combustion is performed using only the diffusion combustion fuel flow path,
A method for operating a pilot combustion burner for a gas turbine, wherein when the gas turbine reaches a predetermined load, premix combustion is performed using only the fuel flow path for premix combustion.

Images