The present invention relates to a gas turbine combustor used for a gas turbine that supplies fuel to compressed high-temperature and high-pressure air to perform combustion and supplies a generated combustion gas to a turbine to obtain rotational power, and to the gas turbine including the gas turbine combustor.
A typical gas turbine includes a compressor, a combustor, and a turbine. Air taken in through an air intake port is compressed by the compressor to be high-temperature and high-pressure compressed air. The combustor supplies fuel to the compressed air and performs combustion, thereby obtaining a high-temperature and high-pressure combustion gas (working fluid). The gas turbine drives the turbine with the combustion gas, thereby driving a generator coupled to the turbine.
In the combustor of the gas turbine having the configuration described above, an inner cylinder is supported in an external cylinder, and a transition piece is coupled to an end of the inner cylinder. The external cylinder, the inner cylinder, and the transition piece constitute a casing. The inner cylinder is provided with a pilot nozzle and a plurality of main fuel nozzles. The external cylinder is provided with a plurality of top hat nozzles on the inner peripheral surface. When an air flow of the compressed air comes into the inner cylinder through an air passage, the fuel is injected from the top hat nozzles. The air-fuel mixture is mixed with the fuel injected from the main fuel nozzles in the inner cylinder to be a swirling flow of a pre-mixture flowing into the transition piece. The air-fuel mixture is also mixed with the fuel injected from the pilot nozzle and ignited and burned. As a result, the air-fuel mixture is turned into a combustion gas, and the combustion gas is injected into the transition piece. As a result, the pre-mixture flowing into the transition piece from the main fuel nozzles is ignited and combusted.
Various gas turbine combustors have been developed, including the combustor described in Patent Literature 1, for example.
Patent Literature 1: Japanese Patent Application Laid-open No. 2005-233574
Before the compressed air flows into the inner cylinder, the gas turbine combustor described above mixes in advance the fuel injected from the top hat nozzles with the compressed air flowing through the air passage. Consequently, the gas turbine combustor can mix the fuel and the air for combustion in the mixture more uniformly in the inner cylinder. The top hat nozzles, however, protrude perpendicularly to the flow of the compressed air. This structure causes separation of the flow around the top hat nozzles, thereby generating areas where the flow of the compressed air is slow downstream. As a result, the flame generated in the inner cylinder and the transition piece may possibly flash back, thereby damaging the top hat nozzles.
To address the disadvantages described above, an object of the present invention is to provide a gas turbine combustor and a gas turbine that suppresses flashback of a flame.
To achieve the object described above, a gas turbine combustor of the present invention is a gas turbine combustor that includes an external cylinder having a tubular shape, an inner cylinder disposed in the external cylinder, a first fuel injecting unit disposed in the inner cylinder, an air passage provided between the external cylinder and the inner cylinder to cause compressed air to flow into the inner cylinder, and a second fuel injecting unit disposed in the air passage. The second fuel injecting unit is disposed at a curve communicating with the inner cylinder downstream in the air passage. The second fuel injecting unit has a base end supported by a curved inner surface of the external cylinder and a distal end extending toward the inner cylinder. The second fuel injecting unit is disposed with a center line in a longitudinal direction inclined to a downstream side or an upstream side in a flow direction of the compressed air flowing through the air passage at a predetermined angle with respect to a perpendicular orthogonal to a tangent to the curved inner surface of the external cylinder.
If the compressed air flows into the air passage, fuel is injected from the second fuel injecting unit to the compressed air to generate a mixture. The mixture flows into the inner cylinder. The fuel is injected from the first fuel injecting unit to the mixture to generate a pre-mixture. The second fuel injecting unit is inclined to the downstream side or the upstream side in the flow direction of the compressed air in the air passage. This structure can suppress separation of the flow around the second fuel injecting unit. Consequently, the present invention can reduce the number of areas where the flow of the compressed air is slow downstream of the second fuel injecting unit, thereby suppressing flashback of a flame.
In the gas turbine combustor, the second fuel injecting unit is disposed with the center line in the longitudinal direction inclined to the downstream side in the flow direction of the compressed air flowing through the air passage with respect to the perpendicular within a range from 10 degrees to 30 degrees.
The second fuel injecting unit is inclined to the downstream side in the flow direction of the compressed air within a range from 10 degrees to 30 degrees. With this structure, the second fuel injecting unit is less likely to be resistance to the flow of the compressed air. Consequently, the present invention can reduce the number of areas where the flow of the compressed air is slow downstream of the second fuel injecting unit.
In the gas turbine combustor, an inner surface of the external cylinder defining the air passage has a first linear portion extending in an axial direction of the external cylinder, a second linear portion extending in a direction orthogonal to the axial direction of the external cylinder, and a curve connecting the first linear portion and the second linear portion, and the base end of the second fuel injecting unit is supported closer to the downstream side in the flow direction of the compressed air flowing through the air passage than an intersection of extensions of the first linear portion and the second linear portion.
Consequently, the present invention can further suppress flashback of a flame. Furthermore, the change in the fuel injection position can improve the combustion stability.
In the gas turbine combustor, the second fuel injecting unit has a cylindrical shape and has a spherical shape on the distal end.
The second fuel injecting unit has a cylindrical shape and has a spherical shape on the distal end. Consequently, the present invention can suppress separation of the compressed air caused by the second fuel injecting unit and reduce the number of areas having slow flow velocity in the wake of the second fuel injecting unit.
A gas turbine of the present invention includes a compressor configured to compress air, a combustor configured to mix compressed air compressed by the compressor with fuel to perform combustion, and a turbine configured to obtain rotational power from a combustion gas generated by the combustor. As the combustor, the gas turbine combustor is used.
In the combustor, the second fuel injecting unit disposed in the air passage provided between the external cylinder and the inner cylinder is inclined to the downstream side or the upstream side in the flow direction of the compressed air at the predetermined angle. This structure can suppress separation of the flow around the second fuel injecting unit. Consequently, the present invention can reduce the number of areas where the flow of the compressed air is slow downstream of the second fuel injecting unit, thereby suppressing flashback of a flame.
In the gas turbine combustor and the gas turbine according to the present invention, the second fuel injecting unit disposed in the air passage provided between the external cylinder and the inner cylinder is inclined to the downstream side or the upstream side in the flow direction of the compressed air at the predetermined angle. Consequently, the present invention can reduce the number of areas where the flow of the compressed air is slow downstream of the second fuel injecting unit, thereby suppressing flashback of a flame.
7 and illustrates an attachment state of a peg in a gas turbine combustor according to the present embodiment.
Exemplary embodiments of a gas turbine combustor and a gas turbine according to the present invention are described below in greater detail with reference to the accompanying drawings. The embodiments are not intended to limit the present invention. If there are a plurality of embodiments, the present invention includes a combination of the embodiments.
As illustrated in
The compressor 11 has an air intake port 20 through which air is taken in. In a compressor casing 21, an inlet guide vane (IGV) 22 is disposed, and a plurality of compressor vanes 23 and compressor blades 24 are alternately disposed in the longitudinal direction (axial direction of a rotor 32, which will be described later). A bleed air chamber 25 is provided outside the compressor casing 21. The combustor 12 supplies fuel to compressed air compressed by the compressor 11 and ignites the fuel and the compressed air, thereby performing combustion. The turbine 13 includes a plurality of turbine vanes 27 and turbine blades 28 alternately disposed in the longitudinal direction (axial direction of the rotor 32, which will be described later) in a turbine casing 26. A flue gas chamber 30 is provided downstream of the turbine casing 26 with a flue gas casing 29 provided therebetween. The flue gas chamber 30 includes a flue gas diffuser 31 connected to the turbine 13.
The rotor (rotating shaft) 32 is provided penetrating the center of the compressor 11, the combustor 12, the turbine 13, and the flue gas chamber 30. The end of the rotor 32 on the compressor 11 side is rotatably supported by a bearing 33, and the end thereof on the flue gas chamber 30 side is rotatably supported by a bearing 34. In the compressor 11, a plurality of disks provided with the respective compressor blades 24 are stacked and fixed on the rotor 32. In the turbine 13, a plurality of disks provided with the respective turbine blades 28 are stacked and fixed on the rotor 32. The end of the rotor 32 on the flue gas chamber 30 side is coupled to a drive shaft of the generator, which is not illustrated.
In the gas turbine 10, the compressor casing 21 of the compressor 11 is supported by a leg 35, the turbine casing 26 of the turbine 13 is supported by a leg 36, and the flue gas chamber 30 is supported by a leg 37.
Air taken in from the air intake port 20 of the compressor 11 passes by the IGV 22, the compressor vanes 23, and compressor blades 24 and is compressed, thereby becoming high-temperature and high-pressure compressed air. The combustor 12 supplies predetermined fuel to the compressed air and performs combustion. The high-temperature and high-pressure combustion gas serving as working fluid generated by the combustor 12 passes by the turbine vanes 27 and the turbine blades 28 of the turbine 13, thereby driving and rotating the rotor 32. As a result, the generator coupled to the rotor 32 is driven. The combustion gas that drives the turbine 13 is released to the atmosphere as a flue gas.
The following describes the combustor (gas turbine combustor) according to the present embodiment in greater detail.
In the combustor 12, as illustrated in
Specifically, as illustrated in
The air passage 51 having a ring shape is formed between the external cylinder 41 and the inner cylinder 42. A plurality of pegs (fuel injecting unit) 58 are provided in the air passage 51. As illustrated in
As illustrated in
If an air flow of high-temperature and high-pressure compressed air comes into the air passage 51, fuel F is injected from the pegs 58 to the compressed air to generate a mixture. The mixture flows into the inner cylinder 42. The mixture flowing into the inner cylinder 42 is mixed with the fuel F injected from the main combustion burners 45 to be a swirling flow of a pre-mixture. The mixture is also mixed with the fuel F injected from the pilot combustion burner 44 and ignited and burned, which is not illustrated. As a result, the mixture is turned into a combustion gas, and the combustion gas is injected into the inner cylinder 42. At this time, part of the combustion gas is injected into the inner cylinder 42 in a manner diffusing around with a flame. As a result, the pre-mixture flowing into the inner cylinder 42 from the main combustion burners 45 is ignited and combusted. In other words, a diffusion flame generated by the pilot fuel F injected from the pilot combustion burner 44 can perform flame holding for stable combustion of lean premixed fuel F supplied from the main combustion burners 45.
The following describes the peg 58 in greater detail.
As illustrated in
As illustrated in
The peg 58 is disposed with its center line O in the longitudinal direction inclined to the downstream side in the flow direction of compressed air A flowing through the air passage 51 at a predetermined angle θ with respect to a perpendicular P orthogonal to a tangent T to the inner surface of the curve 41b of the external cylinder 41. The inclination angle θ is preferably from 10 degrees to 40 degrees and is most preferably approximately 30 degrees.
The base end of the peg 58 is supported closer to the downstream side in the flow direction of the compressed air A flowing through the air passage 51 than the intersection of the extensions of the first linear portion 41a and the second linear portion 41c.
The peg 58 does not necessarily have the shape described above.
As illustrated in
The pegs 58 and 71 do not necessarily have the shape described above. The pegs 58 and 71 may have a polygonal columnar shape or an elliptic cylindrical shape or be tapered, thickened toward the end, or stepped, for example. While the pegs 58 and 71 are disposed with their center line O in the longitudinal direction inclined to the downstream side in the flow direction of the compressed air A with respect to the perpendicular P, they may be disposed with their center line O in the longitudinal direction inclined to the upstream side in the flow direction of the compressed air A with respect to the perpendicular P.
If the compressed air A flows into the air passage 51, the fuel F is injected from the pegs 58 and 71 to the compressed air A to generate a mixture. The mixture flows into the inner cylinder 42 and is mixed with the fuel F injected from the pilot nozzle 53 and the main nozzles 56. The pegs 58 and 71 are inclined to the downstream side in the flow direction of the compressed air A in the air passage 51. This structure can suppress separation of the flow around a second fuel injecting unit. As a result, the number of areas where the flow of the compressed air A is slow downstream of the pegs 58 and 71 is reduced, thereby suppressing flashback of a flame from the inner cylinder 42.
As described above, the gas turbine combustor according to the present embodiment includes the external cylinder 41, the inner cylinder 42, the pilot nozzle 53, the main nozzles 56, the air passage 51, and the pegs 58 and 71. The pilot nozzle 53 and the main nozzles 56 are disposed in the inner cylinder 42. The air passage 51 is provided between the external cylinder 41 and the inner cylinder 42. The pegs 58 and 71 are disposed in the air passage 51. The pegs 58 and 71 are disposed at the curve 41b communicating with the inner cylinder 42 downstream in the air passage 51. The base end of the pegs 58 and 71 is supported by the inner surface of the curve 41b, and the distal end thereof extends toward the inner cylinder 42. The pegs 58 and 71 are disposed with their center line O in the longitudinal direction inclined to the downstream side (or upstream side) in the flow direction of the compressed air A flowing through the air passage 51 at the predetermined angle θ with respect to the perpendicular P orthogonal to the tangent T to the curved surface of the external cylinder 41.
This structure can suppress separation of the flow around the pegs 58 and 71. Consequently, the present embodiment can reduce the number of areas where the flow of the compressed air A is slow downstream of the pegs 58 and 71, thereby suppressing flashback of a flame.
In the gas turbine combustor according to the present embodiment, the pegs 58 and 71 are inclined to the downstream side in the flow direction of the compressed air A within a range from 10 degrees to 30 degrees. With this structure, the pegs 58 and 71 are less likely to be resistance to the flow of the compressed air A. Consequently, the present embodiment can reduce the number of areas where the flow of the compressed air A is slow downstream of the pegs 58 and 71.
In the gas turbine combustor according to the present embodiment, the inner surface of the external cylinder 41 defining the air passage 51 has the first linear portion 41a, the second linear portion 41c, and the curve 41b. The first linear portion 41a extends in the axial direction of the external cylinder 41. The second linear portion 41c extends in a direction orthogonal to the axial direction of the external cylinder 41. The curve 41b connects the first linear portion 41a and the second linear portion 41c. The base end of the pegs 58 and 71 is supported closer to the downstream side in the flow direction of the compressed air A flowing through the air passage 51 than the intersection of the extensions of the first linear portion 41a and the second linear portion 41c. Consequently, the present embodiment can further suppress flashback of a flame. Furthermore, the change in the fuel injection position can improve the combustibility.
In the gas turbine combustor according to the present embodiment, the peg 71 has a cylindrical shape and has the spherical portion 71b on the distal end. With this structure, the peg 71 is less likely to be resistance to the flow of the compressed air A. Consequently, the present embodiment can suppress separation of the compressed air A caused by the peg 71 and reduce the number of areas having slow flow velocity in the wake of the second fuel injecting unit.
The gas turbine according to the present embodiment includes the compressor 11, the combustor 12, and the turbine 13. The compressor 11 compresses air. The combustor 12 mixes the compressed air A compressed by the compressor 11 with the fuel F and performs combustion. The turbine 13 obtains rotational power from the combustion gas generated by the combustor 12. The combustor 12 includes the pegs 58 and 71 disposed at the curve 41b of the air passage 51. The pegs 58 and 71 are inclined to the downstream side (or upstream side) in the flow direction of the compressed air A flowing through the air passage 51 at the predetermined angle θ.
With this structure, the present embodiment can suppress separation of the flow around the pegs 58 and 71. As a result, the amount of compressed air flowing around the pegs 58 and 71 is reduced, thereby reducing the number of areas where the flow of the compressed air A is slow downstream of the pegs 58 and 71. Consequently, the present embodiment can suppress flashback of a flame.
While the pegs 58 and 71 according to the present embodiment are disposed at the curve 41b communicating with the inner cylinder 42 downstream in the air passage 51, they may be disposed at the linear portion 41a or 41b instead of the curve 41b.
10 Gas turbine
11 Compressor
12 Combustor
13 Turbine
41 External cylinder
41
a First linear portion
41
b Curve
41
c Second linear portion
42 Inner cylinder
43 Transition piece
44 Pilot combustion burner
45 Main combustion burner
51 Air passage
53 Pilot nozzle (first fuel injecting unit)
56 Main nozzle (first fuel injecting unit)
58, 71 Peg (second fuel injecting unit)
58
a, 71a Injection hole
71
b Spherical portion
A Compressed air
F Fuel
L1, L2 Length
O Center line
P Perpendicular
T Tangent
θ Inclination angle
Number | Date | Country | Kind |
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2016-043589 | Mar 2016 | JP | national |
Filing Document | Filing Date | Country | Kind |
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PCT/JP2017/008996 | 3/7/2017 | WO | 00 |