KR19990054621A - Combustor of gas turbine - Google Patents

Abstract

Disclosed is a combustor of a gas turbine. The combustor is formed between a cylindrical first liner surrounding at least a portion of the combustion section, a cylindrical second liner arranged concentrically to surround the first liner, and between the first liner and the second liner. An annular passageway through which the outlet is in communication with the combustion unit and an inlet thereof in communication with the compressor of the gas turbine, first fuel supply means for injecting gas fuel into the annular passage, and second fuel supply for injecting liquid fuel into the annular passageway Means; and the second fuel supply means includes injection means for injecting liquid fuel into the annular passage in a spray state, and atomization means for increasing the flow rate of the gas mixture flowing on the annular passage and forming a vortex do. Since the combustor of such a gas turbine can implement uniform spraying characteristics of the liquid fuel in the premixing chamber, there is an advantage that the amount of nitrogen oxides can be reduced.

Description

Combustor of gas turbine

The present invention relates to a combustor of a gas turbine, and more particularly, to a dual fuel combustor of a gas turbine capable of using both liquid fuel and gaseous fuel as a fuel.

In a gas turbine, fuel is combusted with compressed air in one or more combustors, where the compressed air is produced by a compressor. Such combustors typically have a first combustion chamber in which a mixture of fuel and air is combusted in a diffusion combustion process. Fuel is supplied to the first combustion section by a fuel nozzle installed concentrically. In the case of using liquid fuel, the fuel nozzle sprays fuel into combustion air so that the fuel is atomized before the fuel is supplied to the first combustion section. And in order to lower the overall fuel / air concentration ratio, additional air is supplied to the combustion chamber located downstream of the first combustion section. However, since there is a high concentration of fuel / air mixture locally in the first combustion section, the fuel / air mixture is easily ignited at the start of combustion and a good flame stability can be obtained despite the low concentration of fuel / air ratio as a whole. .

As the concentration ratio of fuel / air in the first combustion section increases, the combustion temperature becomes very high, and this high combustion temperature increases the formation of nitrogen oxides during combustion. And these nitrogen oxides cause serious environmental pollution. Low concentrations of fuel / air reduce the formation of nitrogen oxides. However, to form such low concentrations of fuel / air mixtures, the fuel must be widely distributed in compressed air and mixed very well. For this purpose it is essential to premix the fuel with the combustion gas before supplying the fuel / air mixture to the combustion chamber.

In the case of gaseous fuel, gaseous fuel is supplied to the first and second annular passageways to premix the fuel and air, so as to flow into the first combustion section and the second combustion section, respectively. The gaseous fuel supplied to these first and second annular passages is injected through a spray tube installed along the periphery of each passage.

However, such a spray tube is not suitable for atomizing liquid fuel and supplying it to the combustion chamber, such a combustor can be applied only in the case of gaseous fuel.

Therefore, although a spray nozzle for a liquid fuel has been developed, this does not completely atomize the liquid fuel, and thus a problem arises in that a very large fuel droplet is formed in the fuel / air mixture and a locally concentrated fuel / air mixture exists.

The present invention has been made to solve the above problems, and an object thereof is to provide a gas turbine combustor having an improved structure for premixing liquid and fuel so that the atomization characteristics of the liquid fuel are improved.

1 is a schematic configuration diagram of a conventional gas turbine.

FIG. 2 is a cross-sectional view of the combustion section of the gas turbine of FIG. 1.

3 is a cross-sectional view in the longitudinal direction of the combustor of FIG. 2, taken along the line 'B-B' of FIG. 4.

4 is a cross-sectional view taken along the line 'A-A' of FIG.

5 is an enlarged view of a portion 'C' of FIG. 3.

6 is a detailed cross-sectional view of the manifold and pipe of FIG. 4.

Explanation of symbols on the main parts of the drawing

2 ... compressor 4 ... combustor

6 ... turbine 18 ... fuel nozzle

36.First combustion unit 37 ... Second combustion unit

40 ... first liner 42 ... second liner

44 ... 3rd liner 62 ... 1st fuel spray member

76.2nd fuel spray member 72 ... liquid fuel injection pipe

102 Manifold 108 Guide

109..Air distribution hall

The combustor of the gas turbine according to the present invention for achieving the above object, a cylindrical first liner surrounding at least a portion of the combustion unit, and a cylindrical second arranged concentrically to surround the first liner An annular passage formed between the liner, the first liner and the second liner, the outlet of which is in communication with the combustion section, and the inlet of which is in communication with the compressor of the gas turbine; First fuel supply means for mixing gaseous fuel with compressed air to supply the combustion unit through an outlet of the passage, and injecting liquid fuel into the annular passage, mixing the liquid fuel with compressed air to exit the passage. And a second fuel supply means for supplying to the combustion unit through the second fuel supply means, wherein the second fuel supply means includes the annular passageway. And injection means for injecting liquid fuel in a spray state, and atomization means for increasing the flow rate of the gas mixture flowing on the annular passageway and forming a vortex to cause the liquid fuel to be atomized by the compressed air. do.

The injection means includes a ring-shaped manifold provided on one side of the annular passageway, and one or two or more pipes for supplying the liquid fuel to the manifold, and a plurality of fuel injections on the outer surface of the manifold. It is preferable that the holes are formed uniformly along the circumferential direction thereof.

The lower portion of the manifold connected to the pipe is fixed to the outer surface of the second liner, the fuel injection hole is preferably formed in the upper center of the manifold.

A third liner is further provided that surrounds at least a portion of the burner and is surrounded by the second liner, wherein the pipe is located on a second annular passageway formed between the second liner and the third liner.

Preferably, the second annular passage communicates with the compressor to allow cooling compressed air to flow around the pipe.

Preferably, the atomization means has a throat formed by progressively decreasing the cross-sectional area of the downstream portion of the annular passageway than the manifold.

In order to allow the throat to be formed in the annular passage, it is preferable that a flow path guide member is fixed to the first liner or the second liner and extends inclined toward the inside of the passage.

It is preferable that a second air supply means is further provided for supplying compressed air having a velocity component perpendicular to the flow direction of the gas flowing through the throat of the annular passage.

The second air supply means is preferably provided with an air flow hole formed in the throat portion of the first liner.

Hereinafter, with reference to the accompanying drawings, a preferred embodiment of the gas turbine combustor according to the present invention will be described in detail.

1 is a block diagram of a conventional gas turbine.

As shown the gas turbine 1 has a compressor 2 which is connected to a shaft 26 and driven by a turbine 6. The ambient air 12 enters the compressor 2 and is compressed. Compressed air 8 produced in the compressor 2 enters the combustion system, which comprises one or two or more combustors 4, and fuel nozzles 18, which enter both the gaseous fuel 16 and the liquid fuel 14; ). Typically the gaseous fuel 16 is natural gas and the liquid fuel 14 is diesel oil, although other gaseous fuels or liquid fuels may also be used. In the combustor 4 fuel is combusted with the compressed air 8 to produce hot compressed gas 20.

The hot compressed gas 20 produced by the combustor 4 flows into the turbine 6 and expands to generate axial force for driving a load such as a compressor 2 or a generator. The expansion gas 24 generated in the turbine 6 is directly discharged to the outside air or dumped into the outside air after passing through the cogeneration plant.

2 shows the combustion site of the gas turbine.

In the figure only one of the plurality of combustors 4 arranged circumferentially is shown. This combustor 4 is located in a chamber 7 formed by a shell 22, each combustor 4 having a first part 30 and a second part 32. Hot gas 20 is discharged from second portion 32 and enters turbine portion 6 through duct 5. The first part 30 of the combustor is supported by a support plate 28. The support plate 28 extends from the cell 22 and is attached to a cylinder 13 surrounding the first portion 30. The second portion 32 is supported by an arm (not shown) extending from the support plate 28. Thus, by separately supporting the first portion 30 and the second portion 32, it is possible to reduce the thermal stress caused by the difference in thermal expansion.

The combustor 4 has a combustion section composed of a first section and a second section. Referring to FIG. 3, the first combustion section 36 of the combustion section is located inside the first part 30 of the combustor 4, in particular the cylindrical inner liner 44 of the first combustor 30. It is surrounded by a third liner). The inner liner 44 is surrounded by a cylindrical intermediate liner 42 (second liner), and the intermediate liner 42 is surrounded by a cylindrical outer liner 40 (first liner). The three liners 40, 42, 44 are arranged concentrically about an axial center line 71, such that an inner annular passageway 70 (second annular) is formed between the inner and intermediate liners 44, 42. Passageway) and an outer annular passageway 68 (first annular passageway) is formed between the intermediate and outer liner 42, 40.

As shown in FIG. 3, a dual fuel nozzle 18 is located at the inner center of the first portion 30 of the combustor. The fuel nozzle 18 has a cylindrical outer sleeve 48 and an inner sleeve 51 which is also cylindrical in shape. The outer sleeve 48 forms an annular outer passage 56 with a cylindrical intermediate sleeve 49, and the inner sleeve 51 together with the intermediate sleeve 49 annular inner passage 58. To form. A liquid fuel supply tube 60 is located inside the inner sleeve 51 to supply liquid fuel 14 ′ to the liquid fuel spray nozzle 54. The liquid fuel 14 ′ injected from the spray nozzle 54 is injected into the first combustion unit 36 through the fuel outlet 52 formed in the outer sleeve 48. The gaseous fuel 16 ′ flows through the outer passage 56 and is discharged to the first combustion section 36 through the plurality of gaseous fuel ports 50 formed in the outer sleeve 48. Cooling air 38 also flows through the inner passage 58.

In order to supply the fuel / air mixture to the first combustion section 36, it is supplied from the gaseous fuel 16 ″ and the compressor 2 via a first premix passage formed at the front end of the first portion 30 of the combustor. Pre-mixed compressed gas, as shown in Fig. 3. The first premix passage has a first passageway 90 and a second passageway 92 whereby the inlet air flows into two streams 8 '. (8 "). The first passageway 90 has an upstream portion in the radial direction and a downstream portion in the axial direction. The upstream portion is formed between a circular flange 88 extending in the radial direction and a portion extending in the radial direction of the flow path guide 46. The downstream portion is formed between the flow path guide 46 and the outer sleeve 48 of the fuel nozzle 18 and is surrounded by the second passage 92.

The second passage 92 is also composed of an upstream portion in the radial direction and a downstream portion in the axial direction. An upstream portion of the second passage is formed between the radially extending portion of the flow path guide 46 and the radially extending portion of the inner liner 44. And a downstream portion of the second passageway 92 is formed between the axially extending portion of the flow path guide 46 and the axially extending portion of the inner liner 44.

3 and 4, an axially positioned first fuel spray member 62 having a tubular shape is installed around the first premix passage, so that the first and second passages 90 ( Extends transversely through the upstream portion of 92). Gas fuel outlets along the longitudinal direction of the first fuel spray member 62 for introducing gaseous fuel 16 "into the air streams 8 'and 8" flowing through the passages 90 and 92 (64) are distributed. Since the gaseous fuel discharge port 64 has two rows, gaseous fuel 16 "can be injected in the clockwise and counterclockwise directions, respectively.

3 and 4, a plurality of vortex vanes 85,86 are installed upstream of the passages 90,92. The vortex vanes 85 and 86 are preferably installed between the first fuel spray members 62, respectively. As shown in Fig. 4, the vortex vanes 85 and 86 impart counterclockwise and clockwise rotation to the air streams 8 'and 8 ". The imparted vortex improves the mixing of the gaseous fuel 16 "and the air 8 'and 8". This prevents localized large concentrations of mixing in the fuel mixture of the first combustion section 36 from occurring. Thus, generation of nitrogen oxides after combustion is suppressed.

As in FIG. 3, the second burner 37 is formed inside the liner of the second part 32 of the burner 4. The outer annular passage 68 is connected to the second combustion section 37 to form a premixed passage of liquid and gaseous fuel used in the second combustion section 37. A portion 8 "'of compressed air coming from the compressor 2 flows inside the outer annular passage 68.

A plurality of second gaseous fuel spray members 76 provided in the radial direction are installed to penetrate the second premixed passage, that is, the outer annular passage 68 in the transverse direction. The second fuel spray member 76 is supplied with fuel by a circumferentially extending manifold 74. An axially extending fuel supply tube 73 introduces fuel 16 "'into the manifold 74. As shown in Figure 4, there are two rows of spray members 76 along the longitudinal direction. A gaseous fuel outlet 78 is formed to introduce gaseous fuel 16 "'into the second air stream 8"' flowing through the premix passage 68. As shown in FIG. The discharge port 78 discharges the gaseous fuel 16 "'clockwise and counterclockwise, respectively.

According to the invention, the second premix passage ie the outer annular passage 68 is used as a passage for premixing not only the gaseous fuel 16 "'but also the liquid fuel 14" and the compressed air 8 "'. To this end, as shown in Fig. 4, six liquid fuel supply pipes 72 and a ring-shaped manifold 103 connected to the pipes 72 are provided, but the number of the pipes 72 is six. More than 6 or less, each pipe 72 has a first portion extending in the axial direction and a second portion extending in the radial direction, and the first portion is the vortex vane 85. It can be used to support 86. And as shown, the pipe 72 is preferably installed in a second annular passage, ie an inner annular passage 70. Meanwhile, the second annular passage 70 is provided for cooling. Compressed air is intended to flow.

The manifold 102 is formed in a ring shape and is installed in the annular passageway 68. In more detail, the cross-sectional area of the manifold 102 is formed in a semi-circular shape, and the straight portion formed at the lower portion thereof is fixed to the outer surface of the second liner 42 and at the center of the curved portion formed at the upper portion thereof. The fuel injection hole 103 is formed. As shown in FIG. 6, the fuel injection hole 103 is uniformly formed along the circumferential direction of the manifold 102. Meanwhile, six holes 104 are formed in the straight portion of the manifold 102 and the second liner 42 corresponding thereto to connect the six pipes 72 to each other.

The annular passageway 68 forms a throat 69. Thus, the cross-sectional area of the passage 68 in this throat portion is reduced, thereby increasing the velocity of the fluid flow flowing through this portion. This throat 69 is located slightly downstream of the manifold 102, as shown.

Various methods may be used to form the throat 69. As one example, the flow path guide member 108 is installed in the first liner 40. As shown in FIG. The guide member 108 is formed in a predetermined curved shape, one side of which is attached to the upper liner 42, and the other side thereof extends inclined toward the center of the annular passage 68. The inclination direction of the guide member 108 should naturally be in a direction from upstream to downstream of the passageway. By this guide member 108 the cross-sectional area of the passage 68 through which fluid flows is gradually reduced, resulting in a throat 69 at the other end of the guide member 108.

In order to supply separate compressed air to the minimum cross-section of the annular passage 68, that is, the throat 69, an air distribution hole 109 may be formed in the upper liner 40. The air compressed by the compressor 2 supplies the compressed air 110 having the radial velocity component to the throat 108 through the air distribution hole 109. Therefore, the mixture of liquid fuel / air flowing through the annular passage 68 is accelerated in the throat 69 and at the same time a strong vortex is formed by the high pressure air 110 injected from the air distribution hole 109, so that the liquid fuel The atomization is promoted.

Hereinafter, the operation of the combustor of the gas turbine according to the present invention having the configuration as described above will be described.

When gaseous fuel is used, the gaseous fuel 16 'is supplied through the passage 56 of the fuel nozzle 18, and a flame is first generated in the first combustion section 36. As the load on the turbine 6 increases, a higher combustion temperature is required and the additional fuel required is satisfied by the gaseous fuel 16 "supplied through the first fuel spray member 62. First Because the fuel spray member 62 makes the fuel distribution of the compressed air 8 'and 8 "very good, it produces a lower concentration of fuel / air mixture than is supplied from the fuel nozzle 18. This reduces the production of nitrogen oxides in the combustor. Therefore, once ignited in the first combustion unit 36, the gaseous fuel 16 ′ supplied to the fuel nozzle 18 is blocked. And if a fuel amount more than that supplied by the first fuel spray member 62 is required, the additional gaseous fuel 16 "'is supplied to the 2nd combustion part 37 through the 2nd fuel spray member 76. FIG. do.

In the case of using the liquid fuel, the first combustion section 36 is provided by the liquid fuel 14 ′ initially supplied through the liquid fuel supply tube 60 of the fuel nozzle 18 as in the case of the gaseous fuel described above. ) A flame occurs. The additionally required fuel may be satisfied by supplying the liquid fuel 16 "to the second combustion section 37 through the second premix passage 68.

When the liquid fuel 14 "is supplied to the manifold 102 through the pipe 72, this liquid fuel 14" flows clockwise or counterclockwise while passing through the annular passageway through the fuel injection hole 103. 68). The liquid fuel is then mixed by the compressed air 8 "'flowing through the annular passage 68 and atomized to some extent. The liquid fuel / air mixture is flowed downstream of the passage 68, whereby 69. In this case, the flow rate of the fuel / air mixture is drastically increased because the cross section of the passage 69 decreases in the passage 69. In this throat 69, there is also a separate high-pressure air 110. The liquid fuel 14 "becomes completely atomized in this process because) has disturbed the flow flow with velocity components in the direction of flow rate and radial direction. Therefore, since the distribution state of the liquid fuel distributed in the air becomes fairly uniform, the concentration of fuel / air supplied to the second combustion section 37 by the second premixing passage 68 decreases, thereby generating nitrogen oxides. It can be suppressed. Once ignition occurs in the first combustion section 36, the liquid fuel 14 ′ supplied to the fuel nozzle 18 does not need to be increased further, because the demand for additional fuel flow rate is manifold 102. This is because it can be satisfied by the liquid fuel 14 "

Since the pipe 72 for supplying the liquid fuel 14 " to the manifold 102 is located relatively close to the first combustion section 36, a structure for cooling the pipe 72 must be provided. According to the invention, as in Fig. 3, this can be done by forming a plurality of holes 94 in the radially extending portion of the third liner 44. Part of the compressed air supplied from the compressor 2 A pipe 73 located in the passage 70 by causing 66 to flow through the hole 94 into an annular passage 70 formed between the second liner 42 and the third liner 44. Flow around to cool the pipe.

As described above, the combustor of the gas turbine according to the present invention can realize uniform spray characteristics by a low-cost liquid fuel premixing device composed of a manifold, a pipe, and the like, thereby reducing the amount of nitrogen oxide generated in the combustor. There is an advantage that it can.

Although the present invention has been described with reference to one embodiment shown in the drawings, this is merely exemplary, and may be modified in various forms by those skilled in the art within the technical scope of the present invention.

Claims (9)

A cylindrical first liner surrounding at least a portion of the combustion section, A cylindrical second liner disposed concentrically so as to surround the first liner; An annular passage formed between the first liner and the second liner, the outlet of which communicates with the combustion section and whose inlet communicates with the compressor of the gas turbine; A first fuel supply means for injecting gas fuel into the annular passage, mixing the gas fuel with compressed air and supplying the gas fuel to the combustion section through an outlet of the passage; Injecting liquid fuel into the annular passage, comprising: a second fuel supply means for mixing the liquid fuel with compressed air and supplying it to the combustion section through an outlet of the passage; The second fuel supply means, Injection means for injecting liquid fuel in a spray state into the annular passage; And an atomization means for increasing the flow rate of the gas mixture flowing on the annular passageway and forming a vortex, so that the liquid fuel is atomized by the compressed air. The method of claim 1, The injection means, A ring manifold provided on one side of the annular passageway, One or more pipes for supplying the liquid fuel to the manifold; Combustor of the gas turbine, characterized in that a plurality of fuel injection holes are formed uniformly along the circumferential direction on the outer surface of the manifold. The method of claim 2, A lower portion of the manifold connected to the pipe is fixed in contact with an outer surface of the second liner, The fuel injection hole is a combustor of a gas turbine, characterized in that formed in the upper center of the manifold. The method of claim 2, A third liner is further provided that surrounds at least a portion of the combustion unit and is surrounded by the second liner. And the pipe is located on a second annular passageway formed between the second liner and the third liner. The method of claim 4, wherein And the second annular passage communicates with the compressor to allow cooling compressed air to flow around the pipe. The method of claim 1, And said atomizing means comprises a throat formed by progressively decreasing the cross sectional area of the downstream portion of said annular passageway than said manifold. The method of claim 6, In order for the throat to be formed in the annular passageway, And a flow passage guide member fixed to the first liner or the second liner and extending inclined toward the inside of the passage. The method of claim 7, wherein And a second air supply means for supplying compressed air having a velocity component perpendicular to the flow direction of the gas flowing through the throat of the annular passage. The method of claim 8, The second air supply means, Combustor of the gas turbine, characterized in that provided with an air flow hole formed in the throat portion of the first liner.