Patent Application
20150128601
The present application claims priority from Japanese patent application JP 2013-234675 filed on Nov. 13, 2013, the content of which is hereby incorporated by reference into this application.
The present invention relates to a gas turbine combustor.
From a viewpoint of environment protection, the gas turbine combustor is required for a further reduction of the NOx emission. As a measure for reduction of the NOx emission of the gas turbine combustor, a premixing combustor may be cited, though in this case, a flashback is worried that is a phenomenon in which a flame may enter the premixing combustor and damages the combustor.
Japanese Patent Laid-open No. 2003-148734 (Patent Literature 1) discloses a gas turbine combustor which is configured many fuel nozzles for feeding fuel to a combustion chamber and many air holes for feeding air that are positioned on the downstream side of the fuel nozzles and the injection holes of the fuel nozzles and the air holes are arranged coaxially.
Japanese Patent Laid-open No. 2003-148734
The gas turbine combustor is required to be operated stably under wide operation conditions from ignition to full load and reduce the NOx emission.
In the gas turbine combustor disclosed in Patent Literature 1, the multi-burner structure with a plurality of burners arranged and the mixing enhancement structure by fuel nozzles are disclosed, though a problem arises that when combustion air flows in the space wherein a plurality of fuel nozzles are lined on the upstream side of the air hole plates of the burners, a pressure loss due to separating of the flow generated behind the fuel nozzles is caused.
The pressure loss in the gas turbine combustor is related to an efficiency reduction of the entire gas turbine, so that to increase the efficiency of the gas turbine, it is necessary to reduce the pressure loss in the gas turbine combustor.
An object of the present invention is to provide a gas turbine combustor capable of reducing the pressure loss of the gas turbine combustor without increasing the NOx emission.
A gas turbine combustor of the present invention comprising a burner including a plurality of fuel nozzles for injecting fuel, an air hole plates positioned on a downstream side of the fuel nozzles and configured by each of the fuel nozzles and a plurality of air holes arranged in pairs with each of the fuel nozzles, and a combustion chamber for mixing fuel injected from the fuel nozzles configuring the burners and air injected from the air holes and injecting and burning the mixed fuel, characterized in that,
each of the fuel nozzles configuring the burners is provided with a projection in which a part of an outer edge of a section of the fuel nozzle is protruded outward; the projection is arranged so as to be directed toward a center of the gas turbine combustor; and the projection of the fuel nozzle is positioned on a downstream side of a flow of combustion air flowing around each of the fuel nozzles.
According to the present invention, a gas turbine combustor capable of reducing the pressure loss of the gas turbine combustor without increasing the NOx emission can be realized.
The gas turbine combustor which is an embodiment of the present invention will be explained below by referring to the drawings.
The gas turbine combustor which is the first embodiment of the present invention will be explained by referring to
In the gas turbine plant shown in
And, the gas turbine combustor 2 is stored inside a casing 4. Further, the gas turbine combustor 2 includes a burner 6 on the top thereof and an almost cylindrical liner 10 for separating the high-pressure air and the combustion gas inside the combustor 2 on the downstream side of the burner 6.
On the outer periphery of the liner 10, a flow sleeve 11 as an outer peripheral wall forming an air flow path through which the high-pressure air flows down is arranged. The flow sleeve 11 is larger in diameter than the liner 10 and is arranged cylindrically in an almost concentric circle with the liner 10.
Further, on the downstream side of the liner 10, transition piece 12 for leading the high-temperature combustion gas 18 generated in a combustion chamber 5 of the gas turbine combustor 2 is arranged. Further, on the outer periphery side of the transition piece 12, a flow sleeve 13 is arranged.
The suction air 15, after compressed by the compressor 1, becomes the high-pressure air 16 and at the gas turbine rated load, becomes high temperature of 400° C. or higher depending on the pressure ratio.
The high-pressure air 16, after filled in the casing 4, flows into the space between the transition piece 12 and the flow sleeve 13 and cools the transition piece 12 by a convection cooling from the outer wall surface.
Furthermore, the high-pressure air 16, via the circular flow path formed between the flow sleeve 11 and the liner 10, flows toward the top of the gas turbine combustor 2. The high-pressure air 16, in the middle of the flow, is used for the convection cooling of the liner 10.
Further, a part of the high-pressure air 16 is injected from many cooling holes provided in the liner 10 into the liner 10 along the inner wall surface thereof to form a cooling air film and protects and cools the liner 10 from the high-temperature combustion gas 18.
Among the high-pressure air 16, residual combustion air 17 which is not used to cool the liner 10 flows into the combustion chamber 5 from many air holes 32 provided in air hole plates 31 positioned on the wall surface of the combustion chamber 5 on the upstream side.
The combustion air 17 flowing from the many air holes 32 into the liner 10 is burned together with the fuel injected from fuel nozzles 26 in the combustion chamber 5 and generates the high-temperature combustion gas 18.
The high-temperature combustion gas 18 is fed to the turbine 3 via the transition piece 12. The high-temperature combustion gas 18 is discharged after driving the turbine 3 and becomes exhaust gas 19.
The driving force obtained by the turbine 3 is transmitted to the compressor 1 and the generator 8 via the shaft 7. A part of the driving force obtained by the turbine 3 drives the compressor 1, pressurizes air, and generates high-pressure air. Further, another part of the driving force obtained by the turbine 3 rotates the generator 8 to generate electric power.
The burner 6 installed on the top of the gas turbine combustor 2 includes a plurality of fuel systems of fuel systems 51 and 52. The fuel systems 51 and 52 include fuel flow control valves 21 and 22 respectively, and the flow rates of the fuel systems 51 and 52 are adjusted by the fuel flow control valves 21 and 22 respectively, and the power generation rate of a gas turbine plant 9 is controlled.
Further, on the upstream side branching to the plurality of fuel systems 51 and 52, a fuel cutoff valve 20 for cutting off the fuel is installed.
The gas turbine combustor 2 in the present embodiment is configured by one burner 6 and the burner 6 is configured by many fuel nozzles 26, a fuel nozzle header 24 for distributing the fuel to the many fuel nozzles 26, and the air hole plates 31 where the many air holes 32 with air and fuel passing through are arranged in one-to-one correspondence with the fuel nozzles 26.
The fuel nozzles 26 and the air holes 32 formed in the air hole plates 31 are arranged circularly on three rows of concentric circles around a center axis 80 of the burner 6. The combustion air 17 flows in from the outer periphery of the burner 6, by slipping through the gaps of the plurality of fuel nozzles 26 and flowing toward the burner center 80, flows into the air holes 32 formed in the air hole plates 31.
In the air holes 32 of the air hole plates 31, the combustion air 17 and a fuel jet stream 27 are mixed and the mixed gas is fed to the combustion chamber 5. Further, the air holes 32 of the burner are formed so as to be inclined to the axial center of the combustion chamber 5, thus a swirl flow 40 is formed on the downstream side of the burner 6, and by a recirculation flow 41 generated by the swirl flow 40, a flame 42 is formed.
The gas turbine combustor 2 of this embodiment is configured by one burner 6, so that the center axis 80 of the burner 6 and a center axis 81 of the gas turbine combustor 2 coincide with each other.
Here, the shape of the fuel nozzles 26 configuring the burner 6 of the gas turbine combustor 2 in the present embodiment will be shown.
As shown in
Therefore, in the gas turbine combustor 2 of the present embodiment shown in
And, the edge 62 of the projection protruded outside the fuel nozzle 26 is arranged toward the downstream side of the flow of the combustion air 17, thus the flow of the combustion air 17 around the fuel nozzle 26 is adjusted, so that the formation of a recirculation flow due to separating is suppressed and a reduction of the pressure loss of the gas turbine combustor 2 can be realized.
In
As shown in
The edge 62 which is a projection formed at each rear edge of the fuel nozzles 26 configuring the burner 6 of the gas turbine combustor 2 of the present embodiment is arranged so as to be directed to the burner center in the downstream direction of the flow of the combustion air 17.
In
Further, the arrangement of the many air holes 32, if they are arranged circularly in the respective rows, is not restricted to arrangement on a concentric circle with the burner 6 and the center of each circle may be different from the burner center 80.
Further, if the separating of the combustion air flow behind each fuel nozzle 26 can be suppressed, the shape of the section of the fuel nozzle 26 on the upstream side of the flow is not restricted to the round shape as shown in
Further, with respect to the flow in the section shape of the fuel nozzle 26, the shapes of the section of the fuel nozzle 26 on the upstream side and the downstream side, as shown in
To suppress the separating of the flow of the combustion air behind the fuel nozzle 26 and reduce the pressure loss, the shape of the edge 62 in which the rear edge of the fuel nozzle 26 becomes a projection projected outward is optimum, though as shown in
Further, as shown in
In
For the burner 6 of the gas turbine combustor 2 in the present embodiment, the fuel nozzle 26 in the aforementioned structure with the projection formed at the rear edge is used, thus the flow around the fuel nozzle 26 is adjusted and unsteady hydrodynamic force acting on the fuel nozzles 26 caused by the separating of the flow is suppressed and the reliability of the structure of the gas turbine combustor 2 is improved.
Further, on the downstream side of the pairs of the focused fuel nozzle 26 and the air hole 32 formed in the air hole plate 31 to be focused, that is, turbulence of the combustion air 17 flowing into the pairs of the fuel nozzle 26 closer to the center of the burner 6 and the air hole 32 is reduced, so that the flow-in rate of the combustion air into the air hole 32 is unified, and the local fuel air ratio in the combustion chamber 5 of the gas turbine combustor 2 becomes uniform, thus the NOx emission is reduced.
As explained above, according to the present embodiment, a gas turbine combustor capable of reducing the pressure loss without increasing the NOx emission can be realized.
Next, the gas turbine combustor 2 which is the second embodiment of the present invention will be explained by referring to
In the gas turbine combustor 2 of the second embodiment, the explanation of the structure and operation effects common to the gas turbine combustor 2 of the first embodiment is omitted and only the different portions will be explained below.
In the gas turbine combustor 2 of the present embodiment shown in
In the gas turbine combustor 2 of the present embodiment, the structure of the multi-burner 34 as shown in
Even in the multi-burner 34 of the gas turbine combustor 2 shown in the present embodiment, the combustion air 17 flows in from the outer periphery of the multi-burner 34, slips through the gaps of the plurality of fuel nozzles 26 of the outer peripheral burners 36 and the gaps of the plurality of outer peripheral burners 36 and furthermore the gaps of the plurality of fuel nozzles 26 of the central burner 35, flows toward the combustor center 81, and flows into the air holes 32 of the plurality of outer peripheral burners 36 and the central burner 35.
As a fuel nozzle 26 in the gas turbine combustor 2 of the present embodiment, any of the shapes of the fuel nozzle 26 shown in the gas turbine combustor 2 of the first embodiment is acceptable and fuel nozzles in combination of some of the shapes may be installed.
In
In the case of the structure of the multi-burner 34 in the gas turbine combustor 2 of the present embodiment, the center 80 of the central burner 35 of the gas turbine combustor 2 coincides with the center 81 of the gas turbine combustor 2, so that the edge 62 which is the projection at the rear edge of the fuel nozzle 26 is arranged so as to be directed to the center 81 of the burner in the flow direction of the combustion air flow 17.
Namely, it is the same arrangement method as that of the fuel nozzles 26 in the gas turbine combustor 2 of the first embodiment shown in
Therefore, the fuel nozzles 26 of the burner 6 positioned on the outer periphery of the gas turbine combustor 2, as shown in
According to the gas turbine combustor 2 of the present embodiment, similarly to the single burner 6, even in the multi-burner 34, the separating of the flow behind the fuel nozzles 26 is suppressed and the pressure loss can be reduced. In addition, the flow around the fuel nozzles 26 is adjusted, thus the unsteady hydrodynamic force acting on the fuel nozzles 26 caused by the separating of the flow is suppressed and the reliability of the structure of the gas turbine combustor 2 is improved.
Further, on the downstream side of the pairs of the fuel nozzle 26 and the air hole 32 to be focused, that is, turbulence of the combustion air 17 flowing into the pairs of the fuel nozzle 26 and the air hole 32 closer to the combustor center 81 is reduced, so that the flow-in rate of the combustion air 17 into the air hole 32 is unified, and the local fuel air ratio in the combustion chamber 5 of the gas turbine combustor 2 becomes uniform, thus the NOx emission is reduced.
Therefore, according to the present embodiment, even in a gas turbine combustor in which a multi-burner is configured by combining a plurality of burners, the reduction of the pressure loss can be realized without increasing the NOx emission.
As explained above, according to the present embodiment, a gas turbine combustor capable of reducing the pressure loss without increasing the NOx emission can be realized.
Next, the gas turbine combustor 2 which is the third embodiment of the present invention will be explained by referring to
In the gas turbine combustor 2 of the third embodiment shown in
Namely, as for the fuel nozzles 26 arranged in the plurality of circular rows, a fuel nozzle 26 positioned on a more outer periphery side has a larger recirculation flow formed behind it and the pressure loss associated with it is increased.
Therefore, the pressure loss reduction effect due to changing of the shape thereof to the shape of the edge 62 which is the shape of the projection at the rear edge of the fuel nozzle 26 shown in the gas turbine combustor 2 of the first embodiment becomes larger in the fuel nozzle 26 positioned on the outer periphery side than in the fuel nozzle 26 positioned on the inner periphery side.
Meanwhile, in association with the shape change of the projection at the rear edge of each fuel nozzle 26, there are possibilities that the machining costs of the fuel nozzles 26 and the gas turbine combustor itself may increase. To suppress the increase in the machining costs, a method of reducing the number of fuel nozzles 26 whose shape is to be changed may be considered.
In that case, as shown in
Even when the fuel nozzles of the gas turbine combustor 2 are arrayed in four or more circular rows, only the fuel nozzle 26 on the outermost periphery thereof is changed to the exact shape of the edge 62 which is the shape of the projection shown in the fuel nozzle 26 of the gas turbine combustor 2 of the first embodiment, and thereby the effect similar to the case of the fuel nozzles 26, arranged in three rows, of the gas turbine combustor 2 can be obtained.
Further, if the increase in the machining costs is permitted to a certain extent, the shape change of the fuel nozzles 26 is not restricted to the outermost periphery and within the range with the increase permitted, on a priority basis from the outermost periphery, the shape of the fuel nozzles 26 on a plurality of peripheries can be changed.
As mentioned above, according to the gas turbine combustor 2 of the present embodiment, the number of fuel nozzles 26 whose shape is changed is restricted, and thereby the pressure loss reduction can be realized while suppressing the increase in the machining costs.
As explained above, according to the present embodiment, a gas turbine combustor capable of reducing the pressure loss without increasing the NOx emission can be realized.
Next, the gas turbine combustor 2 which is the fourth embodiment of the present invention will be explained by referring to
In the gas turbine combustor 2 of the fourth embodiment shown in
Even in the gas turbine combustor 2 of the present embodiment for forming one multi-burner 34 in combination with a plurality of burners, the flow rate of the combustion air flowing around the fuel nozzles 26 becomes higher as the combustion air is separated from the combustor center 81, so that as the fuel nozzles 26 are separated from the combustor center 81, the recirculation flow formed behind it becomes larger and the pressure loss in association with it also becomes larger. Therefore, the shape thereof is changed to the shape of the fuel nozzles 26 shown in the gas turbine combustor 2 of the first embodiment, and thereby the pressure loss reduction effect becomes higher.
Therefore, a circle 82 having a radius of R with the combustor center 81 as the center is defined and only the fuel nozzles 26 whose centers are positioned outside the circle 82 are changed to the shape of the fuel nozzles 26 shown in the gas turbine combustor 2 of the first embodiment, and thereby the number of nozzles whose shape will be changed is restricted, and by suppressing the increase in the machining costs of the fuel nozzles 26, the pressure loss reduction effect can be maximized.
The radius R of the circle 82 is determined by the changeable number of fuel nozzles which is calculated from the allowable increase in the machining costs or the required magnitude of pressure loss reduction.
As mentioned above, according to the gas turbine combustor 2 of the present embodiment, even in the gas turbine combustor for forming one multi-burner in combination with a plurality of burners, the number of nozzles for changing the shape thereof is restricted, thus the pressure loss reduction can be realized while suppressing the increase in the machining costs.
As explained above, according to the present embodiment, a gas turbine combustor capable of reducing the pressure loss without increasing the NOx emission can be realized.
Next, the gas turbine combustor 2 which is the fifth embodiment of the present invention will be explained by referring to
In the gas turbine combustor 2 of the fifth embodiment shown in
In the gas turbine combustor 2 of the present embodiment, the structure of the fuel nozzle 26 of the gas turbine combustor 2 capable of suppressing the separating of the flow of the combustion air behind the fuel nozzle 26, reducing the pressure loss of the gas turbine combustor, and inserting the tip of the fuel nozzle 26 into the air hole 32 formed in the air plate 31 is shown.
As shown in
However, in the shape of the fuel nozzle 26 of the gas turbine combustor 2 shown in the first embodiment, the maximum width of the section of the fuel nozzle 26 becomes larger than the diameter of the air hole 32 and the fuel nozzle 26 may not be inserted into the air hole 32.
Therefore, in the fuel nozzle 26 of the gas turbine combustor 2 of the present embodiment, as shown in
Further, in the shape of the fuel nozzle 26 of the gas turbine combustor 2 of the present embodiment shown in
Therefore, as shown in
By the aforementioned fuel nozzle 26 of the gas turbine combustor 2 of the present embodiment, the separating of the flow of the combustion air 17 behind the fuel nozzle 26 is suppressed, and the pressure loss of the gas turbine combustor is reduced, and the insertion of the tip of the fuel nozzle 26 into the air hole 32 can be realized.
As explained above, according to the present embodiment, a gas turbine combustor capable of reducing the pressure loss without increasing the NOx emission can be realized.
| Number | Date | Country | Kind |
|---|---|---|---|
| 2013-234675 | Nov 2013 | JP | national |
