专利汇可以提供GAS TURBINE专利检索,专利查询,专利分析的服务。并且In a gas turbine that generates rotational power by supplying fuel to compressed air compressed by a compressor and burning the fuel in a combustor and supplying resultant combustion gas to a turbine, a circumferential distance (S) starting from a leading edge (32c) of a turbine first stage nozzle (32) toward a trailing edge (32d) side of the first stage nozzle (32) and ending at center of transition pieces (22) of such combustors that are adjacent in a circumferential direction is set relative to a circumferential pitch (P) of such first stage nozzles (32) within a range of 0.05≤S/P≤0.15, and an axial distance (L) between a leading edge (32C) of the first stage nozzle (32) and a transition piece rear end (222) of the combustor is set relative to the circumferential pitch (P) of the first stage nozzles (32) within a range of 0.00≤L/P≤0.13. By improving the relative position of a transition piece (33) of the combustor and the first stage nozzle (32), both suppression of inner pressure fluctuations of the combustor and enhancement in aerodynamic efficiency can be achieved.,下面是GAS TURBINE专利的具体信息内容。
The present invention relates to a gas turbine, and more particularly, to a gas turbine with an improved relative position of a combustor transition piece and a turbine first stage nozzle.
A gas turbine includes a compressor, a combustor, and a turbine. The compressor compresses air taken in through an air inlet to make high-temperature, high-pressure compressed air. The combustor supplies fuel to the compressed air and burns the fuel to make high-temperature, high-pressure combustion gas. The turbine is configured to include a plurality of turbine nozzles and turbine rotor blades alternately arranged in a casing. The turbine rotor blades are driven by the combustion gas supplied to an exhaust passage, whereby a rotor connected to a generator is driven to rotate, for example. The combustion gas that has driven the turbine has its pressure converted into static pressure by a diffuser, and is then released into the atmosphere.
Some conventional gas turbines have a carefully devised relative position of a transition piece of the combustor that is an outlet through which the combustion gas is guided toward the turbine and a turbine first stage nozzle that is exposed to the combustion gas first. Such gas turbines are designed to include two (even-numbered multiple) turbine first stage nozzles per combustor, and are so configured that the center of the transition piece of the combustor coincides with the inter-nozzle center at the leading edges of the first stage nozzles. The combustion gas from the combustor is made to pass mainly between the first stage nozzles, thereby lowering the maximum temperature on the surface of the first stage nozzles (see Patent Document 1, for example).
A method is known that enhances turbine efficiency by controlling the relative positional relationship of the transition piece of the combustor and the turbine first stage nozzles (see Patent Document 2, for example). As illustrated in
Depending on the relative position of the combustor and the first stage nozzle, a wake flow (Karman vortex street) developed after the transition piece rear end of the combustor causes edge tones along the leading edge of the turbine first stage nozzle. Resonance of three elements, that is, the frequency of the wake flow, and the frequency and the acoustic eigenvalue of the edge tones, causes inner pressure fluctuations of the combustor, disadvantageously resulting in the occurrence of noise or vibration during its operation. Note that the inner pressure fluctuations mentioned above are distinguishable from inner pressure fluctuations (combustion oscillation) attributable to a combustion state of fuel by their different drive sources. The inner pressure fluctuations that arise from edge tones caused by wake flows are hereinafter simply referred to as the inner pressure fluctuations, unless otherwise specified.
As described above, by placing the transition piece of the combustor and the first stage nozzle closer to each other, the development of wake flows and the inner pressure fluctuations of the combustor caused by the occurrence of edge tones are supposed to be suppressed. However, to enhance turbine efficiency, wake flows need to be flown into the pressure surface side of the first stage nozzle. To this end, the transition piece of the combustor and the first stage nozzle need to be constantly spaced apart by a certain distance, which means suppressing the inner pressure fluctuations and enhancing turbine efficiency are in a trade-off relationship. Patent Document 2 discloses no means to solve them.
The present invention has been made in view of the foregoing, and has an object to provide a gas turbine that can suppress inner pressure fluctuations of a combustor and enhance aerodynamic efficiency.
According to an aspect of the present invention, in a gas turbine that generates rotational power by supplying fuel to compressed air compressed by a compressor and burning the fuel in a combustor and supplying resultant combustion gas to a turbine, a circumferential distance S starting from a leading edge of a turbine first stage nozzle toward a trailing edge side of the first stage nozzle and ending at center of such combustors that are adjacent in a circumferential direction is set relative to a circumferential pitch P of such first stage nozzles within a range of 0.05≤S/P≤0.15, and an axial distance L between a leading edge of the first stage nozzle and a rear end of the combustor is set relative to the circumferential pitch P of the first stage nozzles within a range of 0.00≤L/P≤0.13.
With this gas turbine, the smaller the axial distance L is, the further the development of wake flows after the rear end of the combustor is suppressed. Accordingly, the occurrence of edge tones along the leading edge of the first stage nozzle can be suppressed. In addition, by setting the circumferential distance S relative to the circumferential pitch P within the range of 0.05≤S/P≤0.15, the aerodynamic efficiency of the first stage nozzle can be enhanced in a stable manner.
Advantageously, in the gas turbine, the circumferential distance S is set relative to the circumferential pitch P to satisfy S/P=0.10.
With this gas turbine, the inner pressure fluctuations of the combustor can be further suppressed and the aerodynamic efficiency can be enhanced.
Advantageously, in the gas turbine, the axial distance L is set relative to the circumferential pitch P within a range of 0.08≤L/P≤0.13.
With this gas turbine, even if it is difficult to make the axial distance L relative to the circumferential pitch P satisfy L/P=0, in other words, it is difficult to place the leading edge of the first stage nozzle and the rear end of the combustor closest to each other, the occurrence of edge tones can be suppressed desirably and the inner pressure fluctuations of the combustor can be suppressed.
Advantageously, in the gas turbine, a circumferential thickness D of a rear end of the combustors that are adjacent in the circumferential direction is set relative to the circumferential pitch P within a range of D/P≤0.26.
With this gas turbine thus configured, the occurrence of edge tones can be further suppressed to suppress the inner pressure fluctuations of the combustor, and the aerodynamic efficiency can be enhanced.
According to the present invention, by making the axial distance L smaller, the development of wake flows after the outlet edge of the combustor transition piece can be suppressed, and the occurrence of edge tones along the leading edge of the turbine first stage nozzle can be thus suppressed. Furthermore, by desirably setting the range of the circumferential distance S, the aerodynamic efficiency of the first stage nozzle can be enhanced in a stable manner.
An exemplary embodiment of a gas turbine according to the present invention will now be described in detail with reference to some accompanying drawings. This embodiment is not intended to limit the present invention.
The gas turbine includes, as illustrated in
The compressor 1 compresses air to make compressed air. The compressor 1 includes, in a compressor casing 12 having an air inlet 11 through which air is taken in, a compressor vane 13 and a compressor rotor blade 14. The compressor vane 13 is placed on the compressor casing 12 side, and a plurality of such compressor vanes 13 is provided in the circumferential direction. The compressor rotor blade 14 is placed on the rotor 4 side, and a plurality of such compressor rotor blades 14 is provided in the circumferential direction. The compressor vanes 13 and the compressor rotor blades 14 are arranged alternately along the axial direction.
The combustor 2 supplies fuel to the compressed air compressed by the compressor 1 and ignites the fuel with a burner to make high-temperature, high-pressure combustion gas. The combustor 2 includes an inner cylinder 21 as a combustion cylinder having the burner (not illustrated) and mixing therein the compressed air and the fuel to burn the fuel, a transition piece 22 that guides the combustion gas from the inner cylinder 21 to the turbine 3, and an outer casing 23 that guides the compressed air from the compressor 1 to the inner cylinder 21. A plurality of such combustors 2 is provided in the circumferential direction with respect to a combustor casing 24.
The turbine 3 generates rotational power from the combustion gas combusted by the combustor 2. The turbine 3 includes, in a turbine casing 31, a turbine nozzle 32 and a turbine rotor blade 33. The turbine nozzle 32 is placed on the turbine casing 31 side, and a plurality of such turbine nozzles 32 is provided in the circumferential direction. The turbine rotor blade 33 is placed on the rotor 4 side, and a plurality of such turbine rotor blades 33 is provided in the circumferential direction. The turbine nozzles 32 and the turbine rotor blades 33 are arranged alternately along the axial direction. On the rear side of the turbine casing 31, an exhaust chamber 34 including an exhaust diffuser 34a that communicates with the turbine 3 is provided.
The rotor 4 has one end on the compressor 1 side supported by a bearing 41 and the other end on the exhaust chamber 34 side supported by a bearing 42, and is provided rotatably about the axial center R. The end of the rotor 4 on the exhaust chamber 34 side is connected to a drive shaft of a generator (not illustrated).
In the gas turbine thus configured, the air taken in through the air inlet 11 of the compressor 1 is compressed while passing through the compressor vanes 13 and the compressor rotor blades 14 and turned into high-temperature, high-pressure compressed air. Then, the combustor 2 supplies certain fuel to the compressed air and burns the fuel, whereby high-temperature, high-pressure combustion gas is generated. The combustion gas passes through the turbine nozzles 32 and the turbine rotor blades 33 of the turbine 3, thereby driving the rotor 4 to rotate. By applying rotational power to the generator connected to the rotor 4, electric power is generated. Exhaust gas after driving the rotor 4 to rotate has its pressure converted into static pressure by the exhaust diffuser 34a in the exhaust chamber 34, and is then released into the atmosphere.
In the gas turbine thus configured, the transition piece 22 of the combustor 2 and a turbine first stage nozzle 32 of the turbine 3 that is placed closest to the combustor 2 are placed in the following relationship.
As illustrated in
A circumferential distance S starting from the leading edge 32c (the closest part to the combustor 2 side) of the first stage nozzle 32 toward the trailing edge 32d side of the first stage nozzle 32 and ending at the center of the combustors 2 (the connected transition pieces 22) is set relative to a circumferential pitch P of the first stage_nozzles 32 within the range of 0.05≤S/P≤0.15. In other words, the circumferential distance S is set within the range of equal to or more than 5% and equal to or less than 15% of the circumferential pitch P.
An axial distance L between the leading edge 32c of the first stage nozzle 32 and the transition piece rear end 222 is set relative to the circumferential pitch P of the first stage nozzles 32 within the range of 0.00≤L/P≤0.13. In other words, the axial distance L is set within the range of equal to or more than 0% and equal to or less than 13% of the circumferential pitch P.
A circumferential thickness D of an end of the connected transition pieces 22 of the combustors 2 that are adjacent in the circumferential direction is set relative to the circumferential pitch P within the range of D/P≤0.26. In other words, the circumferential thickness D is set within the range of equal to or less than 26% of the circumferential pitch P.
Analysis results of the present embodiment in which the combustors 2 and the first stage nozzles 32 are placed to satisfy the relationships described above and of comparative examples are plotted in
Referring to
Referring to
As can be apparently seen in
As can be apparently seen in
These analysis results reveal that, as described above, by setting the circumferential distance S relative to the circumferential pitch P within the range of 0.05≤S/P≤0.15 and setting the axial distance L relative to the circumferential pitch P within the range of 0.00≤L/P≤0.13, the occurrence of edge tones can be suppressed to suppress the inner pressure fluctuations of the combustor, and the aerodynamic efficiency can be enhanced.
Furthermore, by setting the circumferential distance S relative to the circumferential pitch P to satisfy S/P=0.10, the occurrence of edge tones can be further suppressed to suppress the inner pressure fluctuations of the combustor, and the aerodynamic efficiency can be enhanced.
When the axial distance L relative to the circumferential pitch P is made to satisfy 0.00=L/P, the resultant configuration is that the leading edge 32c of the first stage nozzle 32 and the transition piece rear end 222 are placed closest to each other. With this configuration, because the development of wake flows after the transition piece rear end 222 of the combustor 2 is suppressed, the occurrence of edge tones can be suppressed to suppress the inner pressure fluctuations of the combustor. In such cases that a seal member is placed between the combustor 2 and the turbine 3, the axial distance L relative to the circumferential pitch P may fail to satisfy 0.00=L/P due to the structural constraints of the gas turbine. In such a case, in consideration of the constraints, the axial distance L is preferably set relative to the circumferential pitch P within the range of 0.08≤L/P≤0.13.
By setting the circumferential thickness D relative to the circumferential pitch P within the range of D/P≤0.26, the occurrence of edge tones can be further suppressed to suppress the inner pressure fluctuations of the combustor, and the aerodynamic efficiency can be enhanced. Making the circumferential thickness D relative to the circumferential pitch P satisfy D/P=0, i.e., D=0, can be achieved by forming the transition pieces 22 of the combustors 2 that are adjacent in the circumferential direction in a single ring shape, for example. If it is difficult to make a configuration that satisfies D/P=0, the circumferential thickness D is preferably set relative to the circumferential pitch P within the range of 0.18≤D/P≤0.26.
As described above, the gas turbine according to the present invention is suitable, with an improved relative position of the combustor transition piece and the turbine first stage nozzle, for achieving both suppression of the inner pressure fluctuations of the combustor and enhancement in the aerodynamic efficiency.
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