JP2012072708A - Gas turbine and method for cooling gas turbine - Google Patents

Abstract

PROBLEM TO BE SOLVED: To provide a gas turbine which suppresses increase in the amount of cooling air used to control a clearance because the high temperature components of the gas turbine are cooled by air extracted from a compressor and air supplied by a blower separately placed whereas power is needed to supply the air extracted from the compressor and the cooling air supplied by the blower and the increase in the number of components to be cooled increases the consumption of the cooling air, resulting in the reduced performance of the gas turbine.SOLUTION: In the gas turbine which comprises a turbine casing cooling path provided to the casing of the gas turbine and an exhaust diffuser cooling path provided to the exhaust diffuser, there is formed at least one path through which a coolant flows from the turbine casing cooling path into the exhaust diffuser cooling path.

Description

  The present invention relates to a gas turbine and a gas turbine cooling method.

  The gas turbine has a configuration in which a rotor (rotary body) is included in a casing (stationary body). A turbine rotor blade is provided on the outer periphery of the rotor, and a gap exists between the tip (outermost periphery side) of the turbine rotor blade and the casing. Leakage loss occurs when the high-temperature and high-pressure mainstream gas passes through this portion, and the performance of the turbine deteriorates. Therefore, in order to improve the performance of the turbine, it is better that the gap between the turbine rotor blade and the shroud is small.

  On the other hand, when the gap between the blade tip and the casing becomes zero, the blade tip and the shroud come into contact with each other, causing destruction. The gap between the blade tip and the casing changes during the operation of the gas turbine due to the thermal expansion and centrifugal elongation of the rotor and casing, etc., so that the gap is damaged due to contact when the casing is assembled (at startup) under all operating conditions. Designed not to be.

  Generally, in industrial gas turbines, this minimum gap appears during startup. This is because the casing has a larger heat capacity and is less likely to warm than the rotor. In this way, when the gap is minimized during start-up, it is necessary to design the gap so that contact does not occur under this start-up condition. Since the gap during rated operation is larger than that during startup, operation is performed with an excessive gap with large leakage loss during rated operation.

  In order to avoid such an excessive gap, some gas turbines use a compressor bleed air or a separate blower air to cool the casing to suppress the thermal expansion of the casing and control the gap. Yes.

  For example, Patent Document 1 discloses a technique in which an impingement cooling manifold is provided on the casing outer surface of a gas turbine to cool the casing.

JP 2008-196490 A

  The hot part of the gas turbine is cooled by the bleed air from the compressor and the air supplied by a separate blower. Power is required to supply cooling air using compressor bleed air or a blower. As the number of objects to be cooled increases, the amount of cooling air consumed also increases, and the performance of the gas turbine decreases.

  For example, in a conventional gas turbine, turbine blades, combustors, exhaust diffusers, and the like are cooled by different air in parallel. When casing cooling is applied, simply adding a casing cooling system to the above configuration increases the amount of cooling air consumption and the power for supplying cooling air. Will be offset by.

  An object of the present invention is to provide a gas turbine in which an increase in the amount of cooling air used for gap control is suppressed.

  A compressor that generates compressed air, a combustor that generates combustion gas by mixing and burning compressed air and fuel generated by the compressor, and a turbine that is rotationally driven by the combustion gas generated by the combustor; A turbine casing for housing the turbine, an exhaust diffuser connected to the turbine casing, an exhaust diffuser cooling passage provided in the exhaust diffuser, and a plurality of turbine casing cooling passages provided in the turbine casing In the gas turbine in which the refrigerant flows through the exhaust diffuser cooling passage and the turbine casing cooling passage, at least one passage through which the refrigerant flows from the turbine casing cooling passage to the exhaust diffuser cooling passage is formed.

  According to the present invention, it is possible to provide a gas turbine capable of cooling the casing with a small increase in the cooling air flow rate.

1 is a partial sectional view of a gas turbine according to a first embodiment of the present invention. The conceptual diagram of the gas turbine to which this invention is applied. The fragmentary sectional view of the gas turbine to which the present invention is applied. The characteristic figure of the rotor blade tip clearance gap of the conventional gas turbine. The characteristic view of the rotor blade tip gap of the gas turbine which is the 1st example of the present invention. The fragmentary sectional view of the gas turbine which is the 2nd example of the present invention. The fragmentary sectional view of the gas turbine which is the 3rd example of the present invention.

  In the invention described in Patent Document 1, an impingement manifold is provided on the outer surface of the casing to cool the casing by impingement. This cooling air is guided from the blower and released to the atmosphere after impingement cooling of the casing, and the total amount of cooling air used by the gas turbine is increased by the amount of casing cooling.

  A high-temperature portion including an exhaust diffuser of a gas turbine is usually cooled by bleed air from a compressor or air by a separate blower. Power is required to supply cooling air from a compressor, blower, or the like. When the casing is cooled to adjust the gap and the amount of cooling air increases, the cooling air supply power also increases. Therefore, a part of the performance improvement by the gap control is offset by this increase in power. For this reason, it is desirable that the casing can be cooled with a smaller increase in the amount of cooling air.

  The gap between the turbine blade and the shroud is desirably as small as possible. However, when the gap becomes zero, the blade and the shroud come into contact with each other, causing damage. For this reason, for example, on the rear stage side of the turbine, the contact is allowed using a combination of a honeycomb seal and a shroud fin to absorb the influence of manufacturing tolerances and the distortion of the casing, and keep the gap small. Since the honeycomb seal cannot be used on the upstream side of the turbine, where the mainstream gas temperature is high, due to heat resistance problems, a margin is provided at the tip of the rotor blade to avoid contact due to manufacturing tolerances and casing distortion. Yes.

  As described above, since the gap is likely to be larger on the front stage side of the gas turbine than the rear stage side, it is desirable that the gap control amount on the front stage side can be increased.

  The air extracted from the compressor is guided to the turbine side through an extraction pipe provided outside the gas turbine. Since an increase in the amount of extraction piping leads to an increase in cost, for example, the extraction stages are limited to several stages, such as an intermediate stage and a subsequent stage of the compressor, and the number of extraction stages is generally smaller than the number of gas turbine stages to be cooled. Although the use of excessively high pressure air results in an increase in loss, the extraction stage is limited, so that there is a portion where cooling air is supplied at a slightly excessive pressure. In order to adjust this excessive pressure to an appropriate pressure, it is conceivable that pressure loss is caused by an orifice or the like. However, it is desirable to avoid pressure adjustment by such an orifice because it wastes pressure.

  First, a schematic configuration of a gas turbine to which the present invention is applied will be described with reference to FIG.

  FIG. 2 is a conceptual diagram of a gas turbine.

  The gas turbine 101 mainly includes a compressor 102, a combustor 103, and a turbine 104. The compressor 102 compresses the atmospheric air 111 to generate compressed air 106, and sends the generated compressed air 106 to the combustor 103. The combustor 103 generates a combustion gas 107 by mixing and burning the compressed air 106 generated by the compressor 102 and fuel, and supplies the combustion gas 107 to the turbine 104.

  The turbine 104 generates a rotational force on the turbine shaft 105 by the combustion gas 107 having high energy of high temperature and pressure supplied from the combustor 103. The device 109 connected to the gas turbine 101 is driven by the rotational force of the turbine shaft 105. After the energy is recovered by the turbine 104, the combustion gas is discharged from the turbine 104 through the exhaust diffuser 113 as exhaust 112.

  A part of the air compressed by the compressor 102 is extracted as turbine cooling air 110 and supplied to the turbine 104 and the exhaust diffuser 113 without passing through the combustor 103.

  FIG. 3 shows a partial cross-sectional view of the gas turbine.

  A turbine 104 shown in FIG. 3 includes a first stage stationary blade 1, a first stage stationary blade 2, a second stage stationary blade 3, a second stage stationary blade 4, a third stage stationary blade 5, and a third stage stationary blade 6. An arrow 8 indicates the flow direction of the combustion gas 107 in the turbine.

  The first stage blade 2 is connected to the outer periphery of the first stage wheel 9, and the second stage wheel 10 to which the second stage blade 4 is connected and the third stage to which the third stage blade 6 is connected. A turbine shaft 105 is configured by stacking with a wheel 11, a compressor rotor 12 connected to the compressor 102, and a spacer 13 together with a stacking bolt. The turbine shaft 105 recovers the energy of the combustion gas 107 discharged from the combustor 103 by the first stage blade 2, the second stage blade 4, and the third stage blade 6, and the compressor 102 and the turbine shaft end. A device 109 such as a generator connected to the unit is driven.

  The turbine shaft 105 is included in the turbine casing 17. The turbine casing 17 is connected with a first stage stationary blade 1, a second stage stationary blade 3, a third stage stationary blade 5, a first stage shroud 14, a second stage shroud 15, and a third stage shroud 16 on the inner peripheral side of the casing. A diaphragm 24 is connected to the inner peripheral side of the second stage stationary blade 3 and the third stage stationary blade 5.

  There are gaps between the first stage blade 2 and the first stage shroud 14, the second stage blade 4 and the second stage shroud 15, the third stage blade 6 and the third stage shroud 16, and the spacer 13 and the diaphragm 24, respectively. The gap is an interface between the stationary body and the rotating body.

  This gap varies depending on the operating condition of the gas turbine. FIG. 4 shows the change trend of the gap. First, immediately after startup, the turbine shaft 105 increases in rotational speed, extends in the radial direction due to centrifugal force, and the gap decreases. Thereafter, the turbine shaft 105, the shrouds 14, 15, 16 and the turbine casing 17 are thermally expanded due to the temperature rise of the mainstream gas. The turbine shaft 105 expands radially outward, and the shrouds 14, 15, 16 expand radially inward to reduce the gap. The turbine casing 17 expands radially outward to widen the gap.

  FIG. 4 shows the characteristics of the blade tip clearance in a conventional gas turbine. In general, the temperature of the turbine shaft 105 and the shrouds 14, 15, and 16 are likely to rise compared to the turbine casing 17. Therefore, before the turbine is thermally stabilized, specifically, the minimum clearance is shown near the time when the rated load is reached. For this reason, also from FIG. 4, the gap during steady operation is larger than the minimum gap.

  The casing cooling structure in the present embodiment will be described with reference to FIG. FIG. 1 is an enlarged view of the turbine casing 17.

  A cooling air header 21 is provided as an annular space in the outer peripheral portion on the front stage side of the turbine casing 17, and an extraction pipe from a compressor is connected to the cooling air header 21. A cooling flow path 22 extending to the rear side in the axial direction is connected to the end face of the cooling air header 21. The cooling flow path 22 has a substantially circular diameter, and is intermittently arranged in the circumferential direction. A substantially circular core plug 18 is inserted into the cooling flow path 22, and the core plug 18 has an impingement hole at least on the radially inner side.

  The cooling air extracted from the compressor is guided to the cooling air header 21 through the extraction piping. The cooling air header is distributed to the inner space of the core plug 18 in the plurality of cooling flow paths 22 provided in the turbine casing 17, and is jetted out as impingement holes from the impingement holes provided in the core plug 18 to collide with the turbine casing 17. Cooling. Thereafter, the cooling air flows down in the cooling flow path 22 toward the rear side in the turbine axial direction. The cooling passage 22 is connected to the exhaust diffuser cooling passage 20 through the communication hole 19, and the cooling air flowing down in the cooling passage 22 is supplied to the exhaust diffuser cooling passage 20 to cool the exhaust diffuser 113.

  FIG. 5 shows the gap characteristics of the gas turbine according to this embodiment. By performing the casing cooling after the gas turbine reaches the rated output, the casing temperature in the steady state decreases, so the temperature difference between the turbine shaft 105 and the turbine casing 17 in the process from the start to the steady state decreases. The difference between the minimum gap and the gap in the steady state is reduced, and the gap in the steady state can be kept smaller than in the conventional case. At this time, since the minimum gap value in the middle of starting can be set to the same level as the conventional one, the performance can be improved without impairing the reliability of the gas turbine. Further, the cooling of the exhaust diffuser 113 during the start-up is supplied by bypassing the casing cooling passage 22 using a valve or the like separately provided with cooling air.

  Moreover, it is difficult to apply a seal structure that can follow the change in the gap on the upstream side of the gas turbine. This is because, in order to apply a labyrinth seal or the like, it is necessary to form a shroud at the blade tip, but when the shroud is formed, the weight of the blade tip surface increases and the stress of the blade becomes excessive. Furthermore, it is difficult to use a seal structure that allows contact with a honeycomb seal or the like due to heat resistance problems.

  Although it is necessary to keep the gap small in order to suppress the leakage loss, the gap must be designed with a margin in order to prevent breakage and the like.

  On the other hand, on the rear stage side, since the temperature of the mainstream gas is low, a honeycomb seal that can allow contact can be applied. Therefore, a design with a small margin is possible, and a gap can be designed small.

  Thus, since the gap on the front stage tends to be excessive, it is desirable that the gap control amount on the front stage can be increased, that is, the casing cooling effect on the front stage can be increased.

  In the structure of this embodiment, impingement cooling is adopted. Impinge cooling shows high cooling performance with a small amount of air, but if the flow from the lateral direction (hereinafter referred to as cross flow) collides with the cooling air that is ejected from the impingement holes and collides with the surface to be cooled, the cooling performance tends to decrease. . In the structure of the present embodiment, the cooling air ejected from the core plug 18 flows in the cooling flow path 22 from the pre-cooling stage side to the rear stage side. Therefore, since the cross flow on the front stage side can be reduced compared to the rear stage side, it is possible to obtain a cooling effect with high impingement cooling on the front stage side. This makes it possible to keep the gap on the front stage side, which conventionally had to be increased, small.

  In general, the cooling effect is higher when the temperature difference between the surface to be cooled and the cooling air is larger. In the present embodiment, a casing cooling flow path 22 is provided inside the turbine casing 17 so that the distance between the surface to be cooled and the inner surface of the casing is reduced. Since the turbine casing 17 has a higher temperature on the inner side in the circumferential direction and a lower temperature on the outer peripheral side, the temperature of the surface to be cooled can be increased by reducing the distance between the surface to be cooled and the inner surface of the casing, thereby cooling the casing more effectively. it can. Therefore, the casing can be cooled with a small amount of air.

  The cooling air of the exhaust diffuser is discharged into the gas turbine exhaust after the exhaust diffuser is cooled. Since the gas turbine exhaust pressure is almost atmospheric pressure, the cooling air for the exhaust diffuser can be low pressure. Moreover, since the use of high-pressure air leads to an increase in compressor power, it is desirable that the cooling air for the exhaust diffuser 113 be as low as possible.

  Further, the cooling air extraction stage from the compressor 102 has about three stages on the front side, the middle side, and the rear side of the compressor. This is because an increase in the number of extraction stages causes an increase in the number of pipes, thereby increasing costs, and is a result of a trade-off between performance and cost.

  Due to such limitation of the number of extraction stages, the supply pressure of the cooling air of the exhaust diffuser 113 is higher than the minimum pressure required for cooling. Since an excessively high pressure leads to an increase in the amount of cooling air, it is general to adjust the amount of cooling air by causing a pressure loss by an orifice or the like. However, it is inefficient to cause a pressure loss only for such flow rate adjustment.

  In the present embodiment, this pressure loss is caused by impingement cooling in the turbine casing 17 so that the pressure is effectively used for promoting cooling. In other words, by eliminating the orifice, a pressure difference necessary for impingement cooling can be secured, so that it is not necessary to increase the cooling air supply pressure.

  Further, in the structure of this embodiment, the cooling flow path 22 and the exhaust diffuser cooling passage 20 are connected by the communication hole 19, and the cooling air that has finished the casing cooling is guided to the exhaust diffuser cooling passage 20 through the communication hole 19. Then, the exhaust diffuser 113 is cooled. In conventional gas turbines, the exhaust diffuser and the casing are cooled with different air, but by reusing the cooling air as in this embodiment, a new increase in the amount of cooling air associated with the application of casing cooling can be achieved. Can be suppressed. Furthermore, if the number of extraction stages from the compressor can be reduced by reusing the cooling air, the cost will be reduced.

  In addition, you may change the shape of an impingement hole, and the cross-sectional area or shape of the core plug 18 or the cooling flow path 22 with a place. For example, the distance between the core plug 18 and the surface to be cooled of the cooling flow path 22 is shortened on the front side where high cooling performance is required, and the cooling performance is improved. The cross-sectional area of the cooling air passage can be increased and the flow velocity of the cross flow can be reduced. Moreover, although the core plug 18 and the cooling flow path 22 have a substantially circular shape, an elliptical shape, a square shape, a triangular shape, or the like may be changed.

  Moreover, although the compressor bleed air was mentioned as an example of a refrigerant | coolant, the refrigerant | coolant may utilize the air and vapor | steam from a blower or another supply source.

  A gas turbine according to the second embodiment will be described with reference to FIG. FIG. 6 is an enlarged view of the turbine casing 17, and turbulent flow promoting ribs 23 are installed in the cooling flow path 22. Other structures are the same as those in the first embodiment.

  In Example 1, impingement cooling was employed as a cooling method, but impingement cooling exhibits a high pressure loss.

  For example, in the combined cycle, the exhaust 112 passes through the heat exchanger. At that time, a pressure loss occurs, so that the exhaust pressure at the turbine outlet increases compared to the simple cycle. This increases the required pressure of the exhaust diffuser cooling air. In addition, it may be difficult to ensure a pressure difference for performing impingement cooling, such as when the extraction stage pressure of cooling air is low.

  In this embodiment, heat transfer is promoted by installing turbulent flow promoting ribs 23 instead of impingement cooling. Since the turbulent flow promoting rib 23 has a lower pressure loss than impingement cooling, cooling with a low pressure difference is possible.

  Various known structures can be applied to the shape of the turbulent flow promoting rib 23. Moreover, you may change the shape of the turbulent flow promotion rib 23 to apply depending on a place. In general, the cooling performance and pressure loss of the turbulent flow promoting rib are proportional to each other. For example, a turbulent flow promoting rib 23 having a high pressure loss is used on the front side that requires high cooling performance, and the pressure on the rear side is pressure. You may employ | adopt the turbulent flow promotion rib 23 of a shape with low loss and cooling performance.

  Further, the shape of the turbulent flow promoting rib 23 does not have to be uniform in all the cooling passages, and the shape may be changed according to the circumferential temperature distribution of the turbine casing 17.

  A gas turbine according to the second embodiment will be described with reference to FIG. FIG. 6 is an enlarged view of the turbine casing 17. A core plug 18 is inserted on the upstream side of the cooling flow path 22, and a turbulent flow promoting rib 23 is installed on the downstream side. Other structures are the same as those in the first embodiment.

  In the structure of the first embodiment, impingement cooling is adopted over the entire length of the cooling flow path 22. In impingement cooling, the cooling performance tends to decrease due to cross flow. Since the front stage side has less cross flow than the rear stage side, it is possible to obtain a cooling effect with high impingement cooling on the front stage side. However, the cooling performance decreases as the influence of the cross flow increases as the rear stage is reached.

  Therefore, in this embodiment, the impingement cooling is performed only on the upstream side, thereby suppressing the number of impingement holes as compared with the first embodiment and increasing the jet speed of the cooling air from the impingement holes, thereby suppressing the influence of the cross flow. , Trying to get high cooling performance. On the other hand, the downstream side where the cross flow becomes large is changed to cooling by the turbulent flow promoting rib.

  By doing in this way, the cooling of the upstream side of the turbine casing 17 in which especially high cooling performance is required can be strengthened more.

  Further, the cooling structure does not have to be uniform in all the cooling passages, and the shape may be changed according to the circumferential temperature distribution of the turbine casing 17. Furthermore, the shape of the impingement hole, the shape of the core plug 18 and the cooling channel 22, and the cross-sectional area may be changed depending on the location.

  According to each embodiment of the present invention described above, it is possible to provide a gas turbine capable of cooling the casing with a small increase in the cooling air flow rate and capable of more preferable clearance control.

DESCRIPTION OF SYMBOLS 1 First-stage stationary blade 2 First-stage stationary blade 3 Second-stage stationary blade 4 Second-stage stationary blade 5 Third-stage stationary blade 6 Third-stage stationary blade 8 Flow direction of main flow gas 9 First-stage wheel 10 Second-stage wheel 11 Third-stage Wheel 12 Compressor rotor 13 Spacer 14 First stage shroud 15 Second stage shroud 16 Third stage shroud 17 Turbine casing 18 Core plug 19 Communication hole 20 Exhaust diffuser cooling passage 21 Cooling air header 22 Cooling passage 23 Turbulence promoting rib 24 Diaphragm DESCRIPTION OF SYMBOLS 101 Gas turbine 102 Compressor 103 Combustor 104 Turbine 105 Turbine shaft 106 Compressed air 107 Combustion gas 110 Turbine cooling air 111 Atmospheric air 112 Exhaust 113 Exhaust diffuser

Claims (8)

A compressor for generating compressed air;
A combustor that generates combustion gas by mixing and burning compressed air and fuel generated by the compressor;
A turbine that is rotationally driven by the combustion gas generated in the combustor;
A turbine casing containing the turbine;
An exhaust diffuser connected to the turbine casing;
An exhaust diffuser cooling passage provided in the exhaust diffuser;
A plurality of turbine casing cooling passages provided in the turbine casing,
In the gas turbine in which the refrigerant flows through the exhaust diffuser cooling passage and the turbine casing cooling passage,
The gas turbine according to claim 1, wherein at least one passage through which the refrigerant flows from the turbine casing cooling passage to the exhaust diffuser cooling passage is formed. A compressor for generating compressed air;
A combustor that generates combustion gas by mixing and burning compressed air and fuel generated by the compressor;
A turbine that is rotationally driven by the combustion gas generated in the combustor;
A turbine casing containing the turbine;
A turbine casing cooling passage provided in the turbine casing,
In the gas turbine in which the refrigerant flows in the turbine casing cooling passage,
A hollow structure is provided in the turbine casing cooling passage, the hollow structure includes a plurality of communication holes that communicate the inner surface and the outer surface, and at least a part of the refrigerant is provided in the hollow structure. And a gas turbine that reaches the turbine casing cooling passage. A compressor for generating compressed air;
A combustor that generates combustion gas by mixing and burning compressed air and fuel generated by the compressor;
A turbine that is rotationally driven by the combustion gas generated in the combustor;
A turbine casing containing the turbine;
A turbine casing cooling passage provided in the turbine casing,
In the gas turbine in which the refrigerant flows in the turbine casing cooling passage,
A gas turbine, wherein a turbulent flow promoting rib is installed in the turbine casing cooling passage. Among the gas turbines according to claim 1,
A hollow structure is provided in the turbine casing cooling passage, the hollow structure includes a plurality of communication holes that communicate the inner surface and the outer surface, and at least a part of the refrigerant is provided in the hollow structure. And a gas turbine that reaches the turbine casing cooling passage. In the gas turbine according to claim 1 or 4,
A gas turbine comprising a turbulent flow promoting rib in the turbine casing cooling passage. The gas turbine according to any one of claims 1 to 5,
A gas turbine characterized in that an annular space into which the refrigerant is introduced is provided upstream of the turbine casing cooling passage. A compressor for generating compressed air;
A combustor that generates combustion gas by mixing and burning compressed air and fuel generated by the compressor;
A turbine that is rotationally driven by the combustion gas generated in the combustor;
A turbine casing containing the turbine;
An exhaust diffuser connected to the turbine casing;
An exhaust diffuser cooling passage provided in the exhaust diffuser;
A gas turbine cooling method comprising a turbine casing cooling passage provided in the turbine casing,
A method for cooling a gas turbine, wherein at least a part of the refrigerant that has flowed through the turbine casing cooling passage is caused to flow into the exhaust diffuser cooling passage. The method for cooling a gas turbine according to claim 7,
A gas turbine cooling method, wherein the turbine casing is impingement cooled by the refrigerant.

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