WO2017090709A1 - ガスタービン、及びその部品温度調節方法 - Google Patents
ガスタービン、及びその部品温度調節方法 Download PDFInfo
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- WO2017090709A1 WO2017090709A1 PCT/JP2016/084916 JP2016084916W WO2017090709A1 WO 2017090709 A1 WO2017090709 A1 WO 2017090709A1 JP 2016084916 W JP2016084916 W JP 2016084916W WO 2017090709 A1 WO2017090709 A1 WO 2017090709A1
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- gas turbine
- component
- air
- turbine
- blade
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C6/00—Plural gas-turbine plants; Combinations of gas-turbine plants with other apparatus; Adaptations of gas- turbine plants for special use
- F02C6/04—Gas-turbine plants providing heated or pressurised working fluid for other apparatus, e.g. without mechanical power output
- F02C6/06—Gas-turbine plants providing heated or pressurised working fluid for other apparatus, e.g. without mechanical power output providing compressed gas
- F02C6/08—Gas-turbine plants providing heated or pressurised working fluid for other apparatus, e.g. without mechanical power output providing compressed gas the gas being bled from the gas-turbine compressor
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
- F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
- F01D11/14—Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing
- F01D11/20—Actively adjusting tip-clearance
- F01D11/24—Actively adjusting tip-clearance by selectively cooling-heating stator or rotor components
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D17/00—Regulating or controlling by varying flow
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
- F01D25/08—Cooling; Heating; Heat-insulation
- F01D25/10—Heating, e.g. warming-up before starting
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
- F01D25/08—Cooling; Heating; Heat-insulation
- F01D25/12—Cooling
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
- F01D25/08—Cooling; Heating; Heat-insulation
- F01D25/14—Casings modified therefor
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/02—Blade-carrying members, e.g. rotors
- F01D5/08—Heating, heat-insulating or cooling means
- F01D5/085—Heating, heat-insulating or cooling means cooling fluid circulating inside the rotor
- F01D5/087—Heating, heat-insulating or cooling means cooling fluid circulating inside the rotor in the radial passages of the rotor disc
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
- F02C7/12—Cooling of plants
- F02C7/16—Cooling of plants characterised by cooling medium
- F02C7/18—Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
- F02C7/12—Cooling of plants
- F02C7/16—Cooling of plants characterised by cooling medium
- F02C7/18—Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
- F02C7/185—Cooling means for reducing the temperature of the cooling air or gas
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
- F02C7/28—Arrangement of seals
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C9/00—Controlling gas-turbine plants; Controlling fuel supply in air- breathing jet-propulsion plants
- F02C9/16—Control of working fluid flow
- F02C9/18—Control of working fluid flow by bleeding, bypassing or acting on variable working fluid interconnections between turbines or compressors or their stages
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D9/00—Stators
- F01D9/06—Fluid supply conduits to nozzles or the like
- F01D9/065—Fluid supply or removal conduits traversing the working fluid flow, e.g. for lubrication-, cooling-, or sealing fluids
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2220/00—Application
- F05D2220/30—Application in turbines
- F05D2220/32—Application in turbines in gas turbines
- F05D2220/321—Application in turbines in gas turbines for a special turbine stage
- F05D2220/3216—Application in turbines in gas turbines for a special turbine stage for a special compressor stage
- F05D2220/3218—Application in turbines in gas turbines for a special turbine stage for a special compressor stage for an intermediate stage of a compressor
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2240/00—Components
- F05D2240/10—Stators
- F05D2240/11—Shroud seal segments
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/20—Heat transfer, e.g. cooling
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/20—Heat transfer, e.g. cooling
- F05D2260/205—Cooling fluid recirculation, i.e. after cooling one or more components is the cooling fluid recovered and used elsewhere for other purposes
Definitions
- the present invention relates to a gas turbine and a component temperature control method thereof.
- This application claims priority based on Japanese Patent Application No. 2015-231053 filed in Japan on November 26, 2015, the contents of which are incorporated herein by reference.
- the gas turbine includes a compressor that compresses air, a combustor that burns fuel in the air compressed by the compressor, and a turbine that is driven by combustion gas from the combustor.
- the turbine includes a turbine rotor that rotates about an axis, a turbine casing that covers the turbine rotor, and a plurality of stationary blade rows that are fixed to the inner peripheral side of the turbine casing.
- the turbine rotor includes a rotor shaft that extends in the axial direction around the axis, and a plurality of blade rows that are fixed to the outer periphery of the rotor shaft and arranged in the axial direction.
- a stationary blade row is arranged on the upstream side in the axial direction of each blade row.
- Each stationary blade row is composed of a plurality of stationary blades arranged in the circumferential direction around the axis.
- Each blade row is composed of a plurality of blades arranged in the circumferential direction around the axis.
- the gap between the inner peripheral surface of the turbine casing and the tip of the rotor blade is generally called tip clearance. Turbine efficiency increases as the tip clearance decreases.
- Patent Document 1 As a method for adjusting the tip clearance, for example, there is a method described in Patent Document 1 below.
- This method is a method of supplying air to a plurality of blade rings constituting a part of the turbine casing and cooling the plurality of blade rings with air during steady operation of the gas turbine.
- a part of the compressed air discharged from the discharge port of the compressor and guided to the combustor is extracted. And after cooling this extracted compressed air with a cooler, it pressurizes with a pressure
- Compressed air from the booster flows into the blade ring on the upstream side in the axial direction among the plurality of blade rings, and then sequentially flows into the blade ring on the downstream side in the axial direction.
- the compressed air flowing out of the blade ring on the downstream side in the axial direction is sent to the combustor to cool the combustor.
- the compressed air that has cooled the combustor is discharged out of the combustor.
- an object of the present invention is to provide a gas turbine capable of effectively utilizing high-temperature and high-pressure compressed air while suppressing facility costs, and a component temperature adjusting method thereof.
- the gas turbine of the first aspect according to the invention for achieving the above object is as follows.
- a compressor having a plurality of compression stages and sequentially compressing air in each of the compression stages; a combustor that generates combustion gas by burning fuel in the compressed air compressed by the compressor; and the combustion gas
- a turbine having a turbine rotor that rotates about the axis thereof, a cylindrical turbine casing that covers the turbine rotor, and a plurality of stationary blade rows that are fixed to the inner peripheral side of the turbine casing, and a plurality of turbines
- An extraction line for extracting compressed air from the intermediate compression stage of the compression stage as extracted air, and leading the extracted air to a first part constituting a part of the turbine casing; and parts constituting the turbine Among them, a second part different from the first part is provided with a part introduction line that guides the bleed air that has passed through the first part, and the first part extends in the axial direction in which the axis extends, Before from the bleed line A first air flow
- a clearance is required between the inner peripheral surface of the turbine casing and the radially outer end of the turbine rotor. This clearance is generally called chip clearance.
- the steady clearance is the tip clearance when the gas turbine is stably operated.
- the steady clearance is large, the flow rate of the combustion gas passing between the radially outer end of the turbine rotor and the inner peripheral surface of the turbine casing increases during steady operation of the gas turbine. For this reason, if the steady clearance is large, the gas turbine performance during steady operation of the gas turbine is lowered. Therefore, in order to improve the gas turbine performance, it is required to reduce this steady clearance.
- the first component can be cooled by introducing the extraction air to the first component via the extraction line. .
- the inner diameter of the turbine casing even if the combustion gas flows through the combustion gas flow path, it is possible to prevent the inner diameter of the turbine casing from increasing due to thermal expansion. Therefore, it is possible to suppress a change in the inner diameter of the turbine casing between the state in which the combustion gas flows through the combustion gas passage and the state in which the combustion gas does not flow through the combustion gas passage. Therefore, in the gas turbine, a change in the tip clearance can be suppressed, and as a result, the steady clearance can be reduced.
- the compressed air from the intermediate compression stage of the compressor can be used as it is for cooling the first component without being cooled by the cooler.
- the second part is a low-pressure part arranged in a pressure environment lower than the pressure of the compressed air at the outlet of the intermediate compression stage. For this reason, in the gas turbine, even if the compressed air from the compressor is not boosted by the booster, the extracted air that has passed through the first component can be flowed to the low pressure component to cool the low pressure component.
- the first part constituting a part of the turbine casing is cooled to reduce the steady clearance, and the low pressure part, which is the second part different from the first part, is cooled, and then the equipment is installed. Cost can be reduced. Furthermore, in the gas turbine, since the first part and the low pressure part are cooled with compressed air from the intermediate compression stage of the compressor, high-temperature and high-pressure compressed air discharged from the compressor through the final compression stage of the compressor is, for example, It can be effectively used as combustion air.
- a gas turbine of a second aspect according to the invention for achieving the above object is as follows:
- the low-pressure component is disposed on the downstream side in the axial direction, which is the side through which the combustion gas flows, with respect to the first air flow path in the first component.
- a gas turbine of a third aspect according to the invention for achieving the above object is as follows:
- the low-pressure component is disposed in the axial direction in a region where the first air flow path in the first component exists.
- a gas turbine of a fourth aspect according to the invention for achieving the above object is as follows:
- the extraction line is connected to an axially downstream end of the first air flow path, and the component introduction line is an axial direction of the first air flow path. Connected to the upstream end.
- the extracted air flows in the first air flow path from the downstream side in the axial direction to the upstream side in the axial direction. That is, in the gas turbine, the flow of the extracted air becomes a counterflow with respect to the flow of the combustion gas. For this reason, in the gas turbine, the first component can be efficiently cooled.
- a gas turbine of a fifth aspect according to the invention for achieving the above object is
- the second air flow path of the low-pressure part can flow out the extracted air from the part introduction line to a combustion gas flow path through which the combustion gas flows. Is formed.
- each of the plurality of stationary blade rows includes a plurality of stationary blades arranged in a circumferential direction with respect to the axis, and the low-pressure component includes the plurality of the low-pressure components.
- the stationary blade rows there are a plurality of stationary blades that constitute the stationary blade row disposed on the downstream side in the axial direction from the first air flow path.
- the stationary blade includes a blade body extending in a radial direction with respect to the axis to form an airfoil, an outer shroud provided radially outside the blade body, and the blade body.
- the second air flow path that flows out from the inner shroud is formed.
- a plurality of stationary blades that are low-pressure parts can be cooled by the extracted air flowing through the second air flow path, and the combustion gas is discharged into the so-called disk cavity by the extracted air flowing out from the second air flow path. It is possible to prevent the combustion gas flowing through the flow path from flowing in.
- the turbine rotor includes a rotor shaft extending in the axial direction with the axis as a center, and a plurality of rotor blade rows arranged in the axial direction with a space therebetween
- the low pressure component constitutes at least one of the plurality of blade rows arranged in a pressure environment lower than the pressure of the compressed air at the outlet of the intermediate compression stage.
- the rotor shaft is formed with a third air flow path extending from an axial end of the rotor shaft to the plurality of blades constituting the low-pressure component, and the component introduction The line includes the third air flow path of the rotor shaft.
- the turbine rotor includes a rotor shaft extending in the axial direction around the axis, and a plurality of moving blade rows arranged in the axial direction at intervals.
- the turbine casing supports the plurality of split rings opposed to the moving blade row in the radial direction with respect to the axis, and supports the plurality of split rings and the plurality of stationary blade rows from the radially outer side.
- a gas turbine according to a tenth aspect of the invention for achieving the above object is
- the blade ring is an integrally formed product with respect to the axial direction in a region where the first air flow path is formed.
- the bleed air that is connected to the bleed line and the component introduction line bypasses the first air flow path, and flows into the bleed line.
- a bypass line that leads to a component introduction line; a steady state in which the extraction air that has flowed into the extraction line flows into the first air flow path; and the bypass line and the component introduction line that have the extraction air that has flowed into the extraction line And a switching device that switches between a bypass state that flows into the low-pressure component.
- the bleed air can flow through the first component and the low pressure component to cool the first component and the low pressure component.
- the switch by setting the switch to the bypass state, the extracted air does not flow through the first component, but flows exclusively into the low-pressure component, so that the second component is cooled without cooling the first component. Can do.
- the state which is cooling the 1st component and the state which does not cool the 1st component are realizable, cooling the 2nd component. Therefore, in the gas turbine, changes in the inner diameter of the turbine casing can be further suppressed.
- a gas turbine of a twelfth aspect according to the invention for achieving the above object is In the gas turbine according to the eleventh aspect, after the start of the turbine, until the output correlation value which is a parameter having a correlation with the output of the gas turbine or the output becomes a first value, A cooling controller is provided that outputs a command to enter a bypass state and outputs a command to the switch to enter the steady state when the output correlation value exceeds the first value.
- a gas turbine of a thirteenth aspect according to the invention for achieving the above object is In the gas turbine according to any one of the first to tenth aspects, heating means for heating the first component, and starting of the turbine rotor at the latest in the start-up process of the turbine, Until the output or the output correlation value, which is a parameter having a correlation with the output, reaches the first value, the heating means heats the first component, and the output correlation value exceeds the first value. And a heating controller for stopping the heating of the first component by the heating means.
- a gas turbine of a fourteenth aspect according to the invention for achieving the above object is
- the heating means is means capable of heating the first component independently of a start / stop operation of the compressor, and the heating controller is a start process of the turbine
- the heating means is caused to heat the first component until a time when the output correlation value reaches the first value from a predetermined time before the time when the turbine rotor starts to accelerate.
- the inner diameter of the turbine casing can be increased when the turbine rotor starts to rotate at a high speed.
- a gas turbine of a fifteenth aspect according to the invention for achieving the above object is
- the heating means is a heater that is provided in the extraction line or the first component and generates heat.
- a gas turbine of a sixteenth aspect according to the invention for achieving the above object is
- the heating means is connected to the extraction line, and a high-temperature air line through which air having a temperature higher than that of the extraction air extracted from the intermediate compression stage flows.
- a switching device wherein the switching device is in a heating state in which air from the high-temperature air line flows into the first air flow path through the extraction line, and the extracted air that has flowed into the extraction line Is switched to a steady state where the air from the high-temperature air line does not flow into the extraction line, and the heating controller controls the operation of the switch.
- a method for adjusting a component temperature of a gas turbine according to a seventeenth aspect of the invention for achieving the above object is as follows: A compressor having a plurality of compression stages and sequentially compressing air in each of the compression stages; a combustor that generates combustion gas by burning fuel in the compressed air compressed by the compressor; and the combustion gas And a turbine having a turbine rotor that rotates about an axis, a cylindrical turbine casing that covers the turbine rotor, and a plurality of stationary blade rows that are fixed to the inner peripheral side of the turbine casing.
- the compressed air is extracted as an extracted air from an intermediate compression stage of the plurality of compression stages, and the extracted air is included in a first part constituting a part of the turbine casing.
- a second air cooling step for flowing the bleed air that has passed through the first component in a second component different from the first component among the components constituting the turbine The second part is the During a low-pressure part being placed under a lower pressure environment than the pressure of the compressed air at the outlet of the compression stage.
- the first component can be cooled by introducing the bleed air to the first component while the combustion gas is flowing in the combustion gas passage in the turbine casing. For this reason, even if the combustion gas flows through the combustion gas flow path, it is possible to prevent the inner diameter of the turbine casing from increasing due to thermal expansion. Therefore, in the component temperature adjustment method, the change in the tip clearance can be suppressed, and as a result, the steady clearance can be reduced.
- the first component is not the compressed air of high temperature and high pressure discharged from the compressor through the final compression stage of the compressor but the compressed air from the intermediate compression stage of the compressor as the extraction air. Flow to parts. For this reason, in the said component temperature control method, compressed air from the intermediate
- the second component for flowing the bleed air that has passed through the first component is a low-pressure component that is disposed in a pressure environment lower than the pressure of the compressed air at the outlet of the intermediate compression stage. For this reason, in the gas turbine, even if the compressed air from the compressor is not boosted by the booster, the extracted air that has passed through the first component can be flowed to the low pressure component to cool the low pressure component.
- the first component constituting a part of the turbine casing is cooled to reduce the steady clearance, and the low pressure component, which is a second component different from the first component, is cooled.
- Equipment costs can be reduced.
- the first component and the low-pressure component are cooled with compressed air from the intermediate compression stage of the compressor, so the high-temperature and high-pressure compressed air discharged from the compressor through the final compression stage of the compressor. Can be effectively used as, for example, combustion air.
- a gas turbine component temperature adjusting method for achieving the above object is as follows.
- the low-pressure component is disposed on the downstream side in the axial direction, which is the side where the combustion gas flows, from the first region where the extracted air flows in the first component. Has been placed.
- a component temperature adjusting method for a gas turbine according to a nineteenth aspect of the invention for achieving the above object is as follows:
- the low-pressure component is arranged in the axial direction within a range where a first region in which the extracted air flows is present in the first component.
- a gas turbine component temperature control method for achieving the above object is as follows: In the gas turbine component temperature adjustment method according to any one of the seventeenth to nineteenth aspects, in the first air cooling step, the extracted air is allowed to flow in the axial direction upstream in the first component.
- a gas turbine component temperature adjusting method according to the twenty-first aspect.
- the extracted air that has flowed through the low-pressure component flows out into a combustion gas flow path through which the combustion gas flows.
- each of the plurality of stationary blade rows includes a plurality of stationary blades arranged in a circumferential direction with respect to the axis.
- the low-pressure component is a plurality of stationary blades constituting at least one stationary blade row disposed in a pressure environment lower than the pressure of the compressed air at the outlet of the intermediate compression stage among the plurality of stationary blade rows. is there.
- the stationary blade includes a blade body extending in a radial direction with respect to the axis to form an airfoil, and an outer side provided on a radially outer side of the blade body.
- a stationary shroud that includes a shroud and an inner shroud provided radially inward of the blade body.
- the air that has passed through the first component is used as the stationary blade that constitutes the low-pressure component. From the outer shroud of the stationary blade and into the stationary blade, through the blade body of the stationary blade and out of the inner shroud of the stationary blade.
- a component temperature control method for a gas turbine according to a twenty-fourth aspect of the invention for achieving the above-described object is as follows:
- the turbine rotor is spaced apart from the rotor shaft extending in the axial direction about the axis.
- a plurality of moving blade rows arranged side by side, and the low pressure component is disposed in a pressure environment lower than the pressure of the compressed air at the outlet of the intermediate compression stage among the plurality of moving blade rows. It is a plurality of moving blades constituting at least one moving blade row.
- the turbine rotor includes a rotor shaft extending in the axial direction around the axis, and spaced apart in the axial direction.
- a plurality of moving blade rows arranged side by side, wherein the turbine casing is located radially outside of the axis, and is opposed to the plurality of moving blade rows in the radial direction, and a plurality of divided rings
- the first part is the blade ring.
- a method for adjusting the temperature of a part of a gas turbine according to a twenty-sixth aspect of the invention for achieving the above object is as follows: In the gas turbine component temperature adjustment method according to any one of the seventeenth to twenty-fifth aspects, when an output correlation value that is a parameter having a correlation with the output of the gas turbine or the output exceeds a first value. The first air cooling step and the second air cooling step are executed.
- a method for adjusting the temperature of parts of a gas turbine according to the twenty-seventh aspect of the invention for achieving the above-described object is as follows:
- the extracted air from the intermediate compression stage is not passed through the first component until the output correlation value becomes the first value after the turbine is started.
- a third air cooling process is performed to flow through the low-pressure parts.
- temperature control method it is possible to suppress the inner diameter reduction of the turbine casing when high-temperature combustion gas does not flow through the combustion gas flow path, or to increase the inner diameter of the turbine casing. Therefore, in the component temperature control method, the steady clearance can be reduced.
- a gas turbine component temperature adjustment method for achieving the above object is as follows:
- the output correlation value is smaller than the first value during the execution of the first air cooling step and the second air cooling step. If it becomes less than the value, the first air cooling step and the second air cooling step are stopped, and a third air cooling step is performed in which the bleed air from the intermediate compression stage flows to the low pressure component without passing through the first component.
- a component temperature control method for a gas turbine according to the twenty-ninth aspect of the invention for achieving the above object is as follows: In the gas turbine component temperature control method according to the 27th or 28th aspect, in the third air cooling step, the extracted air that has passed through the low-pressure component flows out into a combustion gas passage through which the combustion gas flows.
- a gas turbine component temperature adjustment method for achieving the above object, in the gas turbine component temperature adjustment method according to the twenty-sixth aspect, during the start-up of the turbine, from the time when the turbine rotor starts to accelerate at the latest until the output correlation value reaches the first value. The first component is heated, and when the output correlation value exceeds the first value, a heating process for stopping the heating of the first component is performed.
- the first part is heated before the turbine rotor begins to rotate at a high speed, and the inner diameter of the turbine casing is increased, so the gas turbine is hot-started while reducing the steady clearance.
- the gas turbine includes heating means capable of heating the first component independently of the start / stop operation of the compressor, and in the heating step, During the start-up process of the turbine, the heating means is used to heat the first component from a predetermined time before the turbine rotor starts to accelerate until the output correlation value reaches the first value. To heat.
- the inner diameter of the turbine casing can be increased when the turbine rotor starts to rotate at a high speed.
- the gas turbine component temperature control method according to the thirty-second aspect of the invention for achieving the above object is as follows.
- the heating means is a heater that generates heat.
- the gas turbine 1 of the present embodiment combusts a compressor 10 that compresses outside air A to generate compressed air Acom, and burns fuel F from a fuel supply source in the compressed air Aco.
- a combustor 20 that generates gas G and a turbine 30 that is driven by the combustion gas G are provided.
- the compressor 10 includes a compressor rotor 11 that rotates about an axis Ar, and a cylindrical compressor casing 15 that covers the compressor rotor 11.
- the direction in which the axis Ar extends is referred to as an axial direction Da.
- One side of the axial direction Da is defined as an axial upstream side Dau, and the other side of the axial direction Da is defined as an axial downstream side Dad.
- a radial direction with respect to the axis Ar is simply referred to as a radial direction Dr.
- a side away from the axis Ar in the radial direction Dr is defined as a radially outer side Dro
- a side approaching the axis Ar in the radial direction Dr is defined as a radially inner Dri.
- An opening is formed in the axial upstream side Dau portion of the compressor casing 15. This opening forms an air intake port 15i through which the compressor 10 takes in the outside air A from the outside.
- a plurality of stationary blade rows 16 are fixed to the radially inner side Dri of the compressor casing 15. The plurality of stationary blade rows 16 are arranged at intervals in the axial direction Da.
- Each of the plurality of stator blade rows 16 includes a plurality of stator blades 17 arranged in the circumferential direction Dc with respect to the axis Ar.
- the compressor rotor 11 includes a rotor shaft 12 extending in the axial direction Da around the axis line Ar, and a plurality of rotor blade rows 13 fixed to the outer periphery of the rotor shaft 12.
- Each moving blade row 13 is arranged on the upstream side Dau of any one of the stationary blade rows 16.
- Each of the plurality of blade rows 13 is composed of a plurality of blades 14 arranged in the circumferential direction Dc.
- One moving blade row 13 and one stationary blade row 16 adjacent to the axially downstream side Dad of the moving blade row 13 constitute one compression stage 19.
- the compressor 10 of this embodiment is an axial compressor having a plurality of compression stages 19.
- the turbine 30 is disposed on the axially downstream side Dad of the compressor 10.
- the turbine 30 includes a turbine rotor 31 that rotates about an axis Ar and a cylindrical turbine casing 41 that covers the turbine rotor 31.
- a plurality of stationary blade rows 46 are fixed to the radially inner side Dri of the turbine casing 41.
- the plurality of stationary blade rows 46 are arranged at intervals in the axial direction Da.
- Each of the plurality of stationary blade rows 46 includes a plurality of stationary blades 47 arranged in the circumferential direction Dc.
- the turbine rotor 31 includes a rotor shaft 32 extending in the axial direction Da around the axis line Ar, and a plurality of rotor blade rows 33 fixed to the outer periphery of the rotor shaft 32.
- Each moving blade row 33 is disposed on the axially downstream side Dad of any one of the stationary blade rows 46.
- Each of the plurality of moving blade rows 33 includes a plurality of moving blades 34 arranged in
- the gas turbine 1 of the present embodiment further includes an intermediate casing 51, an exhaust chamber 52, and a bearing 55.
- the intermediate casing 51 is disposed between the compressor casing 15 and the turbine casing 41 in the axial direction Da.
- the exhaust chamber 52 is disposed on the axially downstream side Dad of the turbine casing 41.
- the compressor casing 15, the intermediate casing 51, the turbine casing 41, and the exhaust chamber 52 are connected to each other to form the gas turbine casing 3.
- the compressor rotor 11 and the turbine rotor 31 rotate integrally around the same axis Ar.
- the compressor rotor 11 and the turbine rotor 31 constitute the gas turbine rotor 2.
- the gas turbine rotor 2 is supported by bearings 55 at both ends in the axial direction Da.
- a rotor of a generator 9 is connected to the gas turbine rotor 2.
- the combustor 20 is fixed to the intermediate casing 51.
- a fuel line 25 that supplies fuel F to the combustor 20 is connected to the combustor 20.
- the fuel line 25 is provided with a fuel adjustment valve 26 for adjusting the fuel flow rate.
- the gas turbine 1 of the present embodiment further includes an extraction line 61, a component introduction line 66, a bypass line 71, a switch 75, and a controller 100.
- the bleed line 61 bleeds compressed air from the intermediate compression stage 19 a of the plurality of compression stages 19 of the compressor 10 as bleed air and guides it to a first part constituting a part of the turbine casing 41.
- the intermediate compression stage 19a is one of a plurality of compression stages 19 excluding the compression stage 19 on the most upstream side Dau and the compression stage 19 on the most downstream side Dad. Compression stage 19.
- the part introduction line 66 guides the bleed air that has passed through the first part to a second part that is different from the first part among the parts that constitute the turbine 30.
- the bypass line 71 is connected to the extraction line 61 and the component introduction line 66, bypasses the first component, and guides the extraction air to the component introduction line 66.
- the switch 75 has a steady state in which the bleed air that has flowed into the bleed line 61 flows into the first part, and the bleed air that has flowed into the bleed line 61 passes through the bleed line 61, the bypass line 71, and the part introduction line 66 to the second part. It switches between the bypass state to flow into.
- the switching device 75 is composed of a three-way valve. Therefore, hereinafter, this three-way valve may be referred to as a three-way valve 75.
- the three-way valve 75 is provided at a connection position between the extraction line 61 and the bypass line 71.
- the first opening 75a is connected to the line on the compressor 10 side in the extraction line 61, and the second opening 75b is on the line on the turbine 30 side in the extraction line 61. Connected.
- the third opening 75 c is connected to the bypass line 71.
- the switch 75 does not have to be a three-way valve, and can be constituted by two valves, for example.
- the controller 100 includes a fuel control unit 101 that controls the fuel control valve 26 and a component temperature control unit (cooling controller) 102 that controls the switch 75.
- the turbine casing 41 includes a plurality of split rings 42, a plurality of heat shield rings 43, a blade ring 44 as a first part, and a casing body 45.
- the split ring 42 is located on the radially outer side Dro of the moving blade row 33 and faces the moving blade row 33 in the radial direction Dr.
- the blade ring 44 has an annular shape centered on the axis Ar, and is located on the radially outer side Dro of the plurality of split rings 42.
- the heat shield ring 43 is located between the split ring 42 and the stationary blade 47 and the blade ring 44 in the radial direction Dr, and connects the split ring 42, the stationary blade 47 and the blade ring 44.
- the casing main body 45 has an annular shape centered on the axis Ar, and is located on the radially outer side Dro of the blade ring 44.
- the casing body 45 supports the blade ring 44 from the radially outer side Dro.
- An intermediate casing 51 is connected to the axially upstream Dau of the casing main body 45.
- An exhaust chamber 52 is connected to the axially downstream side Dad of the vehicle body 45. In the exhaust chamber 52, an annular inner diffuser 53i and an outer diffuser 53o are arranged around the axis Ar.
- a portion on the downstream side Dad in the axial direction of the turbine rotor 31 is disposed on the radially inner side Dri of the inner diffuser 53i.
- the inner diameter of the outer diffuser 53o is larger than the outer diameter of the inner diffuser 53i.
- the outer diffuser 53o is disposed at an interval on the radially outer side Dro of the inner diffuser 53i.
- An annular space between the radially outer side Dro of the inner diffuser 53i and the radially inner side Dri of the outer diffuser 53o forms a combustion gas flow path 52p.
- the stationary blade 47 includes a wing body 48 extending in the radial direction Dr to form an airfoil, an outer shroud 49 o provided on the radially outer side Dro of the wing body 48, and the diameter of the wing body 48. And an inner shroud 49i provided in the direction inner side Dri. A seal ring 49r is provided on the radially inner side Dri of the inner shroud 49i. Combustion gas G from the combustor 20 flows between the outer shroud 49o and the inner shroud 49i.
- the moving blade 34 includes a blade body 35 that extends in the radial direction Dr to form an airfoil, a platform 36 that is provided on the radially inner side Dri of the blade body 35, and a blade that is provided on the radially inner side Dri of the platform 36. And a root 37.
- a blade root 37 of the moving blade 34 is fitted into the rotor shaft 32.
- the combustion gas G flows between the platform 36 of the moving blade 34 and the split ring 42 located on the radially outer side Dro of the moving blade 34. Therefore, the outer shroud 49o and the inner shroud 49i of the stationary blade 47, the platform 36 of the moving blade 34, and the split ring 42 define a combustion gas passage 41p through which the combustion gas G flows.
- the combustion gas passage 41p has an annular shape centering on the axis Ar.
- the edge of the combustion gas passage 41p on the outer side Dro in the radial direction is defined by the outer shroud 49o of the stationary blade 47 and the split ring 42. Further, the edge of the radially inner side Dri of the combustion gas passage 41p is defined by the inner shroud 49i of the stationary blade 47 and the platform 36 of the moving blade 34.
- This combustion gas passage 41p is connected to the combustion gas passage 52p between the inner diffuser 53i and the outer diffuser 53o described above.
- a blade ring air flow path (first air flow path or first region) 65 extending in the circumferential direction Dc and the axial direction Da is formed. That is, the blade ring air flow path 65 extends in the circumferential direction Dc and also extends in the axial direction Da within the blade ring 44.
- the position in the axial direction Da of the axial upstream end Eu of the blade ring air passage 65 is a position where the first stage stationary blade row 46 is provided in the axial direction Da.
- the position in the axial direction Da of the axial downstream end Ed of the blade ring air flow path 65 is a position between the second stage stationary blade row 46 and the second stage moving blade row 33 in the axial direction Da.
- the first end of the extraction line 61 is connected to the position of the radially outer side Dro in the intermediate compression stage 19 a in the compressor casing 15.
- the second end of the extraction line 61 is connected to the axial downstream end Ed of the blade ring air passage 65.
- the flow path of the bleed line 61 is formed by a flow path in the first bleed pipe 62, a flow path in the second bleed pipe 63, and a flow path in the third bleed pipe 64.
- the first bleed pipe 62 is connected to a position on the radially outer side Dro in the intermediate compression stage 19a in the turbine casing 41, and is connected to a first opening 75a of a three-way valve 75 that is a switch.
- the second bleed pipe 63 is connected to the second opening 75 b of the three-way valve 75 that is a switching device, and is connected to the casing main body 45 of the turbine casing 41.
- the third extraction pipe 64 is connected to the casing body 45 of the turbine casing 41 and is connected to the axial downstream end Ed of the blade ring air passage 65.
- the second extraction pipe 63 and the third extraction pipe 64 communicate with each other.
- the plurality of stationary blades 47 constituting the fourth stage stationary blade row 46a are low-pressure parts.
- a plurality of stationary blades 47 that are low-pressure components are formed with stationary blade air flow paths (second air flow paths) 47p.
- the stationary blade air flow path 47p opens at the surface of the outer shroud 49o of the stationary blade 47 on the radially outer side Dro and opens at the surface of the radially inner side Dri of the inner shroud 49i of the stationary blade 47.
- a first end of a component introduction line 66 is connected to the axial upstream end Eu of the blade ring air passage 65.
- the second end of the component introduction line 66 is connected to the opening of the stationary blade air flow path 47p in the outer shroud 49o of the plurality of stationary blades 47 constituting the fourth stage stationary blade row 46a. For this reason, the compressed air compressed up to the intermediate compression stage 19 a of the compressor 10 passes through the first extraction pipe 62, the bypass line 71 and the component introduction line 66 of the extraction line 61 as a stationary air flow path. It can flow into 47p.
- the flow path of the component introduction line 66 is formed by a flow path in the reusable pipe 67 and a passenger compartment air flow path 68.
- the reusable piping 67 is connected to the axial upstream end Eu of the blade ring air flow path 65 and is connected to the casing main body 45 of the turbine casing 41.
- the bypass line 71 is connected to the reusable piping 67.
- the vehicle interior air flow path 68 is a flow path formed in the vehicle casing main body 45 and the blade ring 44 of the turbine casing 41.
- the passenger compartment air flow path 68 communicates with the reusable pipe 67 and guides the extracted air from the reusable pipe 67 to the stationary blade air flow paths 47p of the plurality of stationary blades 47 constituting the fourth stage stationary blade row 46a.
- outside air A flows into the compressor casing 15 from the air intake port 15 i of the compressor 10.
- the outside air A is sequentially compressed in a plurality of compression stages 19 in the process of flowing in the compressor casing 15 from the axial upstream side Dau to the axial downstream side Dad to become compressed air Acom.
- the compressed air Acom flows into the intermediate casing 51 from the compressor casing 15.
- the combustor 20 is also supplied with fuel F from a fuel supply source. In the combustor 20, the fuel F burns in the compressed air Acom, and high-temperature and high-pressure combustion gas G is generated.
- the high-temperature and high-pressure combustion gas G flows from the combustor 20 into the combustion gas passage 41p of the turbine 30.
- This combustion gas G rotates the turbine rotor 31 in the process of flowing in the combustion gas flow path 41p.
- the temperature of the combustion gas G when flowing from the combustor 20 into the combustion gas flow path 41p of the turbine 30 is as high as several thousand degrees Celsius.
- the temperature of the combustion gas G gradually decreases in the process in which the combustion gas G flows through the combustion gas passage 41p.
- a clearance is required between the radially outer end of the moving blade 34 and the inner peripheral surface of the turbine casing 41 facing the moving blade 34 in the radial direction Dr.
- This clearance is generally called chip clearance CC, and is preferably as small as possible from the viewpoint of turbine performance.
- the mass of the turbine rotor 31, particularly the rotor shaft 32, is larger than the mass of the turbine casing 41.
- the turbine rotor 31 has a larger heat capacity than the turbine casing 41, and the thermal responsiveness to the temperature change of the gas flowing through the combustion gas passage 41 p is lower than that of the turbine casing 41.
- the amount of thermal deformation per unit time is larger than that of the turbine rotor 31 that is not exposed to the outside air when the gas turbine is stopped. Therefore, when the temperature of the gas flowing through the combustion gas passage 41p changes, the tip clearance CC changes due to the difference in thermal responsiveness between the turbine rotor 31 and the turbine casing 41. In particular, the change in the tip clearance CC when the gas turbine 1 is started and stopped is large.
- the steady clearance is the tip clearance when the gas turbine 1 continues to operate stably and both the turbine rotor 31 and the turbine casing 41 are continuously at the same temperature.
- this steady clearance is large, the flow rate of the combustion gas G passing between the radially outer end of the rotor blade 34 and the inner peripheral surface of the turbine casing 41 increases during steady operation of the gas turbine 1. For this reason, if the steady clearance is large, the performance of the gas turbine during steady operation of the gas turbine 1 is degraded.
- the temperatures of the turbine rotor 31 and the turbine casing 41 are lowered.
- the turbine casing 41 has a smaller heat capacity than the turbine rotor 31 and is exposed to the outside air as described above, the temperature of the turbine casing 41 is abrupt compared to the temperature of the turbine rotor 31. To drop. For this reason, the temperature of the turbine rotor 31 is temporarily higher than the temperature of the turbine casing 41. In this state, the tip clearance CC is small.
- the outer diameter of the turbine rotor 31 is increased by the centrifugal force acting on the turbine rotor 31, and the tip clearance CC is further decreased. That is, when the gas turbine 1 is hot-started, the tip clearance CC becomes extremely small. Further, the radially outer end of the moving blade 34 and the inner peripheral surface of the turbine casing 41 may come into contact with each other. Therefore, even when the gas turbine 1 is hot-started, it is necessary to determine a steady clearance so as to ensure the tip clearance CC.
- the blade ring 44 which is one of the components constituting the turbine casing 41, is cooled from the bleed air under a certain condition, and the change in the inner diameter of the blade ring 44 is reduced, thereby reducing the tip clearance CC.
- the change is made small.
- the three-way valve 75 which is a switching device, communicates with the first opening 75a and the third opening 75c, and does not communicate with the first opening 75a and the second opening 75b. That is, the switch 75 is in a bypass state.
- the gas turbine rotor 2 starts to increase in speed, and the rotation speed N of the gas turbine rotor 2 is lower than a rated rotation speed Nr (for example, 3600 rpm).
- Fuel supply to the combustor 20 is started.
- the high-temperature and high-pressure combustion gas G is generated, and this combustion gas G flows into the combustion gas passage 41p.
- the bleed air that is compressed air from the intermediate compression stage 19 a passes through the stationary blade air flow path 47 p formed in the stationary blade 47 and enters the disk cavity 32 c between the stationary blade 47 and the rotor shaft 32. After flowing in, it passes between the inner shroud 49i of the stationary blade 47 constituting the fourth stage stationary blade row 46a and the platform 36 of the rotor blade 34 constituting the third stage moving blade row 33, and enters the combustion gas flow path 41p. It flows out (S1: 3rd air cooling process).
- a space between the inner shroud 49i of the stationary blade 47 constituting the fourth stage stationary blade row 46a and the platform 36 of the rotor blade 34 constituting the third stage moving blade row 33 is sealed by this bleed air, and the combustion gas flow from there The combustion gas G flowing through the path 41p does not flow into the disk cavity 32c.
- the generator 9 connected to the gas turbine rotor 2 and an external power system are connected.
- the gas turbine output PWx that is, the power from the generator 9 starts to be supplied to the power system.
- the electric power output from the generator 9 is measured with a wattmeter.
- the fuel control unit 101 of the controller 100 receives the power measured by the power meter, that is, the gas turbine output PWx, as well as an external load command and the like.
- the fuel control unit 101 determines the flow rate of fuel supplied to the combustor 20 based on the gas turbine output PWx, the load command, and the like.
- the fuel control unit 101 outputs a valve opening degree command to the fuel control valve 26 based on the fuel flow rate.
- the component temperature control unit 102 of the controller 100 receives the gas turbine output PWx from the fuel control unit 101 and controls the switch 75 based on the gas turbine output PWx.
- the component temperature control unit 102 determines whether or not the gas turbine output PWx exceeds a predetermined first value PW1 (S2). If the gas turbine output PWx does not exceed the predetermined first value PW1, the component temperature control unit 102 causes the switch 75 to maintain the bypass state. That is, the component temperature control unit 102 continues the third air cooling step (S1). On the other hand, when the gas turbine output PWx exceeds the first value PW1, the component temperature control unit 102 outputs a steady state command to the switch 75, and places the switch 75 in a steady state as shown in FIG.
- the bleed air that has passed through the blade ring air flow path 65 is guided to the plurality of stationary blades 47 constituting the fourth stage stationary blade row 46 a via the component introduction line 66.
- This bleed air passes through the stationary blade air flow path 47p formed in the stationary blade 47 and enters the disk cavity 32c between the stationary blade 47 and the rotor shaft 32, as in the third air cooling step (S1) described above. After flowing in, it passes between the inner shroud 49i of the stationary blade 47 constituting the fourth stage stationary blade row 46a and the platform 36 of the rotor blade 34 constituting the third stage moving blade row 33, and enters the combustion gas flow path 41p.
- the plurality of stationary blades 47 constituting the fourth stage stationary blade row 46a is cooled by the extracted air. Further, the space between the inner shroud 49i of the stationary blade 47 constituting the fourth stage stationary blade row 46a and the platform 36 of the moving blade 34 constituting the third stage rotor blade row 33 is sealed by the extracted air.
- the pressure of the extraction air which is the compressed air compressed up to the intermediate compression stage 19a is, for example, 4 kPa. Further, the pressure between the third stage moving blade row 33 and the fourth stage stationary blade row 46a in the combustion gas passage 41p is, for example, 2 kPa. That is, the pressure at the location where the extracted air flows out is lower than the pressure of the extracted air that is the compressed air compressed up to the intermediate compression stage 19a.
- the component temperature control unit 102 determines whether or not the gas turbine output PWx is greater than 0 and less than a predetermined second value PW2 (S5). If the gas turbine output PWx does not become less than the second value PW2, the component temperature control unit 102 causes the switch 75 to maintain a steady state. That is, the component temperature control unit 102 continues the first air cooling step (S3) and the second air cooling step (S4). On the other hand, when the gas turbine output PWx is greater than 0 and less than the second value PW2, the component temperature control unit 102 outputs a bypass state command to the switch 75, and the switch 75 is bypassed as shown in FIG. To.
- the first opening 75a and the third opening 75c communicate with each other, and the first opening 75a and the second opening 75b do not communicate with each other.
- a part of the compressed air compressed up to the intermediate compression stage 19a constitutes the fourth stage stationary blade row 46a through the extraction line 61, the switch 75, the bypass line 71, and the component introduction line 66 as the extraction air.
- the second value PW2 is a value smaller than the first value PW1.
- the reason why the second value PW2 smaller than the first value PW1 is used as the threshold value for switching the switch 75 from the steady state to the bypass state is to prevent the switch 75 from hunting.
- the component temperature control unit 102 determines whether the gas turbine output PWx is greater than 0 and less than the second value PW2, exceeds the first value PW1, or becomes 0 (S7).
- the component temperature control unit 102 causes the switch 75 to maintain the bypass state. That is, the component temperature control unit 102 continues the third air cooling step (S6).
- the component temperature control unit 102 outputs a steady state command to the switch 75, and as shown in FIG. A 1st air cooling process (S3) and a 2nd air cooling process (S4) are performed.
- the component temperature control unit 102 ends the control of the switch 75.
- the switch 75 maintains the bypass state even after the control of the switch 75 by the component temperature control unit 102 is completed. For this reason, before the start-up of the gas turbine 1, the switching device 75 is in a bypass state as described above.
- the blade ring 44 which is the first component, is provided via the extraction line 61.
- the blade ring 44 can be cooled by guiding the bleed air to For this reason, even if the combustion gas flows in the combustion gas flow path 41p and the gas turbine output PWx exceeds the first value PW1, in this embodiment, the inner diameter of the turbine casing 41 is suppressed from increasing due to thermal expansion. can do. Therefore, in this embodiment, the combustion gas G flows through the combustion gas passage 41p and the gas turbine output PWx exceeds the first value PW1, and the combustion gas G does not flow through the combustion gas passage 41p. The change in the inner diameter of the turbine casing 41 can be suppressed.
- the change in the tip clearance CC can be suppressed, and as a result, the steady clearance can be reduced while ensuring the tip clearance CC when the gas turbine 1 is hot-started.
- the first component is not the high-temperature and high-pressure compressed air discharged from the compressor 10 through the final compression stage 19 of the compressor 10 but the compressed air from the intermediate compression stage 19a of the compressor 10 as the extraction air. It is made to flow to the wing ring 44. Therefore, in the present embodiment, the compressed air from the intermediate compression stage 19a of the compressor 10 can be used as it is for cooling the blade ring 44 without being cooled by the cooler.
- the low-pressure component (second component) in the present embodiment is a stationary blade 47 constituting the fourth-stage stationary blade row 46a disposed on the downstream side Dad in the axial direction from the first air flow path of the first component. That is, the low-pressure component (second component) in the present embodiment is the stationary blade 47 arranged in a pressure environment lower than the pressure of the compressed air at the outlet of the intermediate compression stage 19a. For this reason, in the present embodiment, even if the compressed air from the compressor 10 is not boosted by the booster, the bleed air that has passed through the blade ring 44 that is the first part flows to the stationary blade 47 that is the low-pressure part, The stationary blade 47 can be cooled.
- the first part constituting a part of the turbine casing 41 is cooled to reduce the steady clearance and the low pressure part, which is the second part different from the first part, Equipment costs can be reduced. Furthermore, in this embodiment, since the first component and the low-pressure component are cooled by the compressed air from the intermediate compression stage 19a of the compressor 10, the high-temperature and high-pressure discharged from the compressor 10 through the final compression stage 19 of the compressor 10 is used.
- the compressed air Acom can be effectively used as, for example, combustion air.
- the gas turbine output PWx is used as a threshold parameter for switching the state of the switch 75.
- any parameter may be used as long as it has a correlation with the gas turbine output.
- the flow rate of fuel supplied to the combustor 20, the temperature of the combustion gas G at the inlet of the combustion gas passage 41p, and the like may be used as parameters.
- the second end of the bleed line 61 is connected to the axial downstream end Ed of the blade ring air flow path 65, and components are introduced to the axial upstream end Eu of the blade ring air flow path 65.
- the first end of the line 66 is connected.
- the second end of the extraction line 61 a is connected to the axial upstream end Eu of the blade ring air flow path 65.
- a first end of the component introduction line 66 a is connected to the downstream end Ed in the axial direction of the annular air flow path 65.
- the extraction air flows in the blade ring air flow path 65 from the axial downstream side Dad to the axial upstream side Dau.
- the gas turbine 1a of this embodiment is different from the gas turbine 1 of 1st embodiment in the above point, and the other point is the same as the gas turbine 1 of 1st embodiment.
- the second end of the extraction line 61a is connected to the axial upstream end Eu of the blade ring air flow path 65, the axial direction between the first end and the second end of the extraction line 61a.
- the distance of Da can be shortened.
- the first end of the component introduction line 66a is connected to the axial downstream end Ed of the blade ring air flow path 65, the first end and the second end of the component introduction line 66a are connected to each other.
- the distance in the axial direction Da can be shortened. For this reason, in this embodiment, since the length of each line 61a, 66a can be shortened, an installation cost can be restrained somewhat rather than 1st embodiment.
- the temperature change of the combustion gas G in the combustion gas passage 41p and the temperature change of the extracted air in the blade ring air passage 65 will be described with reference to FIG.
- the horizontal axis in FIG. 7 shows the axial direction position on the basis of the inlet of the combustion gas flow path 41p, and a vertical axis
- shaft shows temperature.
- the temperature of the combustion gas G gradually decreases as it flows through the combustion gas passage 41p to the downstream side Dad in the axial direction. For this reason, the temperature of the blade ring 44 that constitutes a part of the turbine casing 41 gradually decreases as the temperature of the combustion gas G flows through the combustion gas passage 41p to the downstream side Dad in the axial direction.
- the extracted air flows in the blade ring air flow path 65 from the axial downstream side Dad to the axial upstream side Dau. That is, in the gas turbine 1 according to the first embodiment, the flow of the extracted air is opposed to the flow of the combustion gas G.
- the temperature of the bleed air in the first embodiment gradually increases due to heat exchange with the blade ring 44 as it flows in the blade ring air flow path 65 from the axial downstream side Dad to the axial upstream side Dau.
- the temperature change amount of the combustion gas and the blade ring 44 accompanying the position change is larger than the temperature change of the extracted air of the first embodiment accompanying the position change.
- the temperature difference Ti1 between the temperature of the extracted air at the position of the downstream end Ed in the axial direction of the blade ring air flow path 65 and the temperature of the combustion gas G and the blade ring 44 at the same position in the axial direction Da The temperature difference To1 between the temperature of the extracted air at the position of the upstream end Eu in the axial direction of the blade ring air flow path 65 and the temperature of the combustion gas and the blade ring 44 at the same position in the axial direction Da becomes larger. For this reason, in the first embodiment, the amount of heat exchange between the extracted air and the blade ring 44 increases as the extracted air flows through the blade ring air flow path 65.
- the extracted air flows in the blade ring air flow path 65 from the axial upstream side Dau to the axial downstream side Dad.
- the temperature of the extraction air in the present embodiment rises due to heat exchange with the blade ring 44 as it flows in the blade ring air flow path 65 from the axial upstream side Dau to the axial downstream side Dad.
- the combustion gas G flows to the axial downstream side Dad and gradually decreases in temperature as it flows to the axial downstream side Dad.
- a temperature difference To2 between the temperature of the extracted air at the position of the downstream end Ed in the axial direction of the ring air flow path 65 and the temperature of the combustion gas and the blade ring 44 at the same position in the axial direction Da becomes smaller. For this reason, in the second embodiment, the amount of heat exchange between the extracted air and the blade ring 44 decreases as the extracted air flows through the blade ring air flow path 65.
- the difference To2 between the temperature of the extracted air at the position of the downstream end Ed in the axial direction of the blade ring air flow path 65 in the second embodiment and the temperature of the combustion gas and the blade ring 44 at the same position in the axial direction Da is The temperature difference Ti1, To1, Ti2 described above is the smallest.
- the position where the extracted air flows into the blade ring air flow path 65 of the second embodiment is the position where the temperature of the combustion gas G and the blade ring 44 is the highest, and the temperature of the extracted air is low.
- the temperature of the extraction air that has flowed into the annular air flow path 65 rises rapidly. Thereafter, the temperature of the bleed air gradually rises as it flows to the axially downstream side Dad.
- the temperature increase rate of the extraction air after the rapid temperature increase immediately after flowing into the blade ring air flow path 65 is lower than the temperature increase rate of the extraction air flowing through the blade ring air flow path 65 in the first embodiment. .
- the temperature of the bleed air not only in the axial downstream end Ed of the blade ring air flow path 65 in the second embodiment but also in each position in a wide region including the axial downstream end Ed is the same in the axial direction Da.
- the difference between the combustion gas at the position and the temperature of the blade ring 44 is also reduced.
- the temperature difference between the temperature of the bleed air and the temperature of the blade ring 44 at each position in the entire area in the axial direction Da of the blade ring air flow path 65 is the blade ring air flow in the second embodiment.
- the temperature difference between the temperature of the extracted air and the temperature of the blade ring 44 at each position in the region including the axially downstream end Ed of the path 65 becomes larger.
- the amount of heat exchange between the bleed air and the blade ring 44 is changed over the entire area of the axial direction Da of the blade ring air flow path 65 in the axial direction of the blade ring air flow path 65 in the second embodiment. It can be higher than the region including the downstream end Ed. Therefore, in the first embodiment, the blade ring 44 can be efficiently cooled in the entire region in the axial direction Da of the blade ring air flow path 65.
- the aspect of the second embodiment it is preferable to employ the aspect of the second embodiment from the viewpoint of suppressing the equipment cost.
- the aspect of the first embodiment it is preferable to employ the aspect of the first embodiment. For this reason, it is preferable to determine which embodiment is adopted from the two viewpoints of equipment cost and efficient cooling of the blade ring 44.
- the position in the axial direction Da of the axial upstream end Eu of the blade ring air passage 65 is the position where the first stage stationary blade row 46 is provided in the axial direction Da. .
- the position in the axial direction Da of the axial downstream end Ed of the blade ring air flow path 65 is a position between the second stage stationary blade row 46 and the second stage moving blade row 33 in the axial direction Da.
- the position in the axial direction Da of the axial upstream end Eu of the blade ring air flow path 65b is the first stage in the axial direction Da as in the first embodiment. This is the position where the stationary blade row 46 is provided.
- the position in the axial direction Da of the axial downstream end Ed of the blade ring air flow path 65 is a position between the third stage stationary blade row 46 and the third stage moving blade row 33 in the axial direction Da. That is, in the gas turbine 1b of this embodiment, the position of the axial downstream end Ed of the blade ring air flow path 65b is set to the axis of the axial downstream end Ed of the blade ring air flow path 65 of the first embodiment.
- the length in the axial direction Da of the blade ring air flow path 65b of the present embodiment is longer than the length of the blade ring air flow path 65 in the first embodiment in the axial direction Da.
- the gas turbine 1b of this embodiment differs from the gas turbine 1 of 1st embodiment in the above points, and the other points are the same as the gas turbine 1 of 1st embodiment.
- the position in the axial direction Da of the axial downstream end Ed of the blade ring air flow path 65b is the axial direction of the axial downstream end Ed of the blade ring air flow path 65 of the first embodiment. Since the position is shifted to the axial downstream Dad from the position of Da, the axial downstream Dad portion of the blade ring 44 can be cooled more than the first embodiment.
- the second end of the extraction line 61 is connected to the axial upstream end Eu of the blade ring air flow path 65, and the blade ring air flow path 65 is axially downstream.
- the first end of the component introduction line 66 may be connected to the end Ed.
- the blade ring of this embodiment and other embodiments may be composed of a plurality of parts.
- the blade ring air flow path may be formed over a plurality of parts constituting the blade ring.
- the blade ring is preferably an integrally formed product with respect to the axial direction Da in the region where the blade ring air flow path is formed from the viewpoint of equipment cost and the like.
- the integrally formed product can be divided with respect to the circumferential direction Dc from the viewpoint of disassembling the gas turbine and the like.
- the gas turbine 1c of this embodiment includes an extraction line 61 and a component introduction line 66, as in the first embodiment. However, the gas turbine 1c of this embodiment does not have the bypass line 71 and the switch 75 in the first embodiment. Further, the gas turbine 1 c of the present embodiment includes a heater (heating unit) 80 that is provided in the extraction line 61 and heats the extraction line 61.
- the heater 80 includes, for example, an electric heater 81 provided along the extraction line 61 and a heater drive circuit 82 that drives the electric heater 81.
- the heater 80 is controlled by a command from a component temperature control unit (cooling controller, heating controller) 102c of the controller 100c.
- a command from the component temperature control unit 102 c is input to the heater drive circuit 82.
- the component temperature control unit 102c determines whether or not it is X hours before the start of the acceleration of the turbine rotor 31 in the starting process of the turbine 30 (S10). The component temperature control unit 102c makes this determination according to information from the outside or from the fuel control unit 101. When the component temperature control unit 102c determines that X hours before the start of the acceleration of the turbine rotor 31 has been reached, the component temperature control unit 102c outputs a heating state command to the heater 80 as shown in FIG. . That is, the component temperature control unit 102 c heats the extraction line 61 by the heater 80.
- the gas turbine rotor 2 generally rotates at a low speed in order to suppress deformation of the gas turbine rotor 2 even after the supply of fuel to the combustor 20 is stopped. For this reason, in the gas turbine casing 3, even when the fuel supply to the combustor 20 is stopped, there is a gas flow similar to that when the fuel F is supplied to the combustor 20. However, the speed of this gas flow is extremely smaller than when the fuel is supplied to the combustor 20. For this reason, even before the turbine 30 is started, the extraction line 61 has an air flow flowing toward the blade ring air flow path 65. Therefore, even before the turbine 30 is started, when the extraction line 61 is heated by the heater 80, the air is heated in the extraction line 61 and then flows into the blade ring air flow path 65. 44 is heated (S11: heating step).
- This extracted air passes through the stationary blade air flow path 47p formed in the stationary blade 47 and flows into the disk cavity 32c between the stationary blade 47 and the rotor shaft 32, and then constitutes the fourth stage stationary blade row 46a. And flows between the inner shroud 49i of the stationary blade 47 and the platform 36 of the moving blade 34 constituting the third-stage moving blade row 33, and flows out into the combustion gas flow path 41p. For this reason, the inner shroud 49i of the stationary blade 47 constituting the fourth stage stationary blade row 46a and the platform 36 of the moving blade 34 constituting the third stage stationary blade row 33 are sealed by the extracted air. However, since the extracted air is heated by the heater 80, it does not function as cooling air for the stationary blades 47 constituting the fourth stage stationary blade row 46a as in the first embodiment.
- the component temperature control unit 102c determines whether or not the gas turbine output PWx exceeds a predetermined first value PW1 (S12). If the gas turbine output PWx does not exceed the first value PW1, the component temperature control unit 102c causes the heater 80 to maintain a heating state. That is, the component temperature control unit 102c continues the heating process (S10). For this reason, in the starting process of the gas turbine 1c, until the gas turbine output PWx exceeds the first value PW1, the inner diameter of the blade ring 44 becomes larger than the case where the extracted air is not heated.
- the component temperature control unit 102c outputs a steady state command to the heater 80, and the heater 80 is in a non-heated steady state as shown in FIG. Put it in a state.
- the heater 80 is in a steady state, the extracted air that is not heated by the heater 80 flows through the blade ring air flow path 65.
- the extracted air flowing through the blade ring air flow path 65 exchanges heat with the blade ring 44 to cool the blade ring 44 (S13: first air cooling step). For this reason, when the gas turbine output PWx exceeds the first value PW1 and the temperature of the combustion gas G at the inlet of the combustion gas passage 41p is high, an increase in the inner diameter of the blade ring 44 due to thermal expansion is suppressed.
- the bleed air that has passed through the blade ring air flow path 65 is guided to the plurality of stationary blades 47 constituting the fourth stage stationary blade row 46 a via the component introduction line 66.
- This extracted air passes through the stationary blade air flow path 47p formed in the stationary blade 47 and flows into the disk cavity 32c between the stationary blade 47 and the rotor shaft 32, and then constitutes the fourth stage stationary blade row 46a.
- the inner shroud 49i of the stationary blade 47 and the platform 36 of the moving blade 34 constituting the third stage moving blade row 33 flow out to the combustion gas passage 41p (S14: second air cooling step). For this reason, the plurality of stationary blades 47 constituting the fourth stage stationary blade row 46a is cooled by the extracted air. Further, the space between the inner shroud 49i of the stationary blade 47 constituting the fourth stage stationary blade row 46a and the platform 36 of the moving blade 34 constituting the third stage rotor blade row 33 is sealed by the extracted air.
- the component temperature control unit 102c determines whether or not the gas turbine output PWx has become 0 (S15). If the gas turbine output PWx is not 0, the component temperature control unit 102c causes the heater 80 to maintain a steady state. That is, the component temperature control unit 102c continues the first air cooling step (S13) and the second air cooling step (S14). On the other hand, if the gas turbine output PWx is 0, the component temperature control unit 102c ends the control of the switch 75.
- the blade ring 44 as the first component is cooled. can do.
- the inner diameter of the turbine casing 41 is increased due to thermal expansion. Can be suppressed. Therefore, also in the present embodiment, the change in the inner diameter of the turbine casing 41 between the state after the combustion gas G flows through the combustion gas passage 41p and the gas turbine output PWx exceeds the first value PW1 and the state before that. Can be suppressed.
- the tip clearance CC is maintained even when the gas turbine 1c is hot-started. It can be suppressed to 0 or extremely small.
- the steady clearance can be made smaller than that in the first embodiment while ensuring the tip clearance CC when the gas turbine 1c is hot-started.
- a cooler that cools the air compressed by the compressor 10 and a booster that pressurizes the air are unnecessary, so that the equipment cost can be reduced. Furthermore, also in this embodiment, since the first component and the low pressure component are cooled by the compressed air from the intermediate compression stage 19a of the compressor 10, the high temperature and high pressure discharged from the compressor 10 through the final compression stage 19 of the compressor 10 is used.
- the compressed air can be effectively used as, for example, combustion air.
- the heater 80 having the electric heater 81 is used.
- any heater that generates heat may be used as the heater, for example, from steam or the gas turbine 1c. You may use as a heater which uses exhaust gas as a heat source.
- the heater 80 is provided in the extraction line 61, you may provide the heater 80 in the blade ring 44 which is 1st components.
- this embodiment is a modification of said 1st embodiment, you may provide a heater similarly to this embodiment also in said 2nd embodiment and said 3rd embodiment.
- the gas turbine 1d of the present embodiment includes an extraction line 61 (hereinafter referred to as the first extraction line 61) and a component introduction line 66, as in the first embodiment. However, the gas turbine 1d of this embodiment does not have the bypass line 71 and the switch 75 of the first embodiment. Further, the gas turbine 1 d of the present embodiment includes a second extraction line (high temperature air line) 84 and a switch 85. In the present embodiment, the second extraction line 84 and the switch 85 constitute a heating unit that heats the blade ring 44.
- the first end of the second extraction line 84 is located in the compressor casing 15 at the position of the high-pressure compression stage 19b on the downstream side of the position where the first end of the first extraction line 61 is connected, or in the intermediate casing 51. It is connected to the. For this reason, compressed air having a higher temperature and pressure than the compressed air flowing into the first bleed line 61 flows into the second bleed line 84.
- a second end of the second extraction line 84 is connected to the first extraction line 61.
- the switch 85 of this embodiment is a three-way valve. The switch 85 is provided at a connection position between the first bleed line 61 and the second bleed line 84.
- the switch 85 includes a steady state in which the first bleed air that is the compressed air that has flowed into the first bleed line 61 flows into the blade ring air flow path 65, and a second state that is the compressed air that has flowed into the second bleed line 84. Switching between the heated state in which the extracted air flows into the blade ring air flow path 65 is performed.
- the switch 85 is controlled by a command from the component temperature control unit (cooling controller, heating controller) 102d of the controller 100d.
- this switch 85 does not need to be a three-way valve, For example, it can also be comprised with two valves.
- the heating state command when the heating state command is output from the component temperature control unit 102d to the switch 85 and the switch 85 is in a heated state, the high-temperature second extracted air from the second extraction line 84 is converted into the blade ring air flow path. Flows into 65. For this reason, when a heating state command is output from the component temperature control unit 102d to the switch 85, a heating process for heating the blade ring 44 is executed as in the fourth embodiment.
- the steady state command when the steady state command is output from the component temperature control unit 102d to the switch 85 and the switch 85 is in a steady state, the low-temperature first extraction air from the first extraction line 61 is the blade ring air. It flows into the flow path 65. For this reason, when a steady-state command is output from the component temperature control unit 102d to the switch 85, a cooling process for cooling the blade ring 44 is executed as in the fourth embodiment.
- the switch 85 of the present embodiment is in a heated state until the gas turbine output PWx exceeds the first value PW1 in the starting process of the gas turbine 1d, similarly to the heater 80 of the fourth embodiment. Moreover, the switch 85 of this embodiment will be in a steady state after the gas turbine output PWx exceeds 1st value PW1 similarly to the heater 80 of the said 4th embodiment.
- the steady clearance can be made smaller than in the first embodiment while ensuring the tip clearance CC when the gas turbine 1d is hot-started.
- the switch 85 is in a heated state from X hours before the start of the acceleration of the turbine rotor 31 in the start-up process of the turbine 30, as long as the compressor rotor 11 does not rotate, the high-temperature second extraction Since air does not flow through the second extraction line 84, the heating step is not performed.
- high-temperature air is generated independently from the start-stop operation of the compressor 10.
- the air from the high-temperature air generation source may be guided to the extraction line 61 by a high-temperature air line.
- 2nd extraction line 84 (high temperature air line) is similar to this embodiment.
- a switch 85 may be provided.
- the component introduction line 66 in the first embodiment has a first end connected to the axial upstream end Eu of the blade ring air flow path 65, and a second end of the plurality of stationary blades constituting the fourth stage stationary blade row 46a. 47.
- the component introduction line 66e in the gas turbine 1e of the present embodiment has its first end connected to the axial upstream end Eu of the blade ring air passage 65 as in the first embodiment.
- the ends are connected to a plurality of moving blades 34 constituting the fourth stage moving blade row 33a. That is, in the present embodiment, the plurality of moving blades 34 constituting the fourth stage moving blade row 33a constitutes a low pressure component.
- a moving blade air flow path (second air flow path) 34p is formed in the moving blade 34 constituting the fourth stage moving blade row 33a.
- the blade air flow path 34p opens at the surface of the blade root radial inner side Dri, and extends from here to the blade body 35 via the blade root 37 and the platform 36. Therefore, the air that has flowed into the rotor blade air flow path 34p flows out to the combustion gas flow path 41p.
- the moving blade air flow path 34 p in the wing body 35 is branched at a plurality of openings and opens on the surface of the wing body 35.
- a second end of the component introduction line 66e is connected to an opening on the blade root 37 side in the moving blade air flow path 34p.
- the flow path of the component introduction line 66e in this embodiment is formed by a flow path in the reusable pipe 67e and a rotor air flow path (third air flow path) 69.
- the reusable pipe 67 e is connected to the axial upstream end Eu of the blade ring air flow path 65 and is connected to the axial downstream end of the turbine rotor 31.
- the rotor air flow channel 69 opens at the downstream end of the turbine rotor 31, and branches from the main flow channel 69a on the axis Ar toward the axial upstream side Dau, the main flow channel 69a, and the fourth stage moving blade row 33a. And a plurality of branch flow paths 69b extending to the blade roots 37 of the plurality of moving blades 34 constituting the same.
- the flow path in the reusable pipe 67 e communicates with the main flow path 69 a of the rotor air flow path 69.
- the reusable piping 67e and the axially downstream end of the rotating turbine rotor 31 are not in contact with each other. Therefore, the fact that the reusable pipe 67e and the axially downstream end of the turbine rotor 31 are connected means that the air from the reusable pipe 67e can flow into the rotor air passage 69 of the turbine rotor 31. Means that.
- the plurality of moving blades 34 constituting the fourth-stage moving blade row 33a can be cooled from the extracted air flowing into the component introduction line 66e.
- a plurality of moving blades 34 constituting the fourth stage moving blade row 33a are provided. It may be a low-pressure part.
- the position in the axial direction Da of the axial upstream end Eu of the blade ring air passage 65 is the position where the first stage stationary blade row 46 is provided in the axial direction Da. .
- the position in the axial direction Da of the axial downstream end Ed of the blade ring air flow path 65 is a position between the second stage stationary blade row 46 and the second stage moving blade row 33 in the axial direction Da.
- the position in the axial direction Da of the axial upstream end Eu of the blade ring air flow path 65f is the first stage in the axial direction Da as in the first embodiment. This is the position where the stationary blade row 46 is provided.
- the position in the axial direction Da of the axial downstream end Ed of the blade ring air flow path 65f is a position where the fourth stage moving blade row 33a is provided in the axial direction Da. That is, in the gas turbine 1f of the present embodiment, the position in the axial direction Da of the axial downstream end Ed of the blade ring air flow path 65f is set to the position of the blade ring air flow paths 65 and 65b of the first embodiment and the third embodiment.
- the length in the axial direction Da of the blade ring air flow path 65f of the present embodiment is longer than the length of the blade ring air flow path 65, 65b in the first embodiment and the third embodiment in the axial direction Da.
- the gas turbine 1f of the present embodiment is different from the gas turbine 1 of the first embodiment in the above points, and the other points are the same as the gas turbine 1 of the first embodiment.
- the position in the axial direction Da of the axial downstream end Ed of the blade ring air flow path 65f is the axis of the blade ring air flow paths 65 and 65b of the first embodiment and the third embodiment. Since the position of the axial downstream end Ed is shifted to the axial downstream side Dad with respect to the axial downstream end Ed, the portion of the blade ring 44 on the axial downstream side Dad can be cooled more than the first embodiment and the third embodiment. it can.
- the position in the axial direction Da of the axial downstream end Ed of the blade ring air flow path 65f is the position where the fourth stage moving blade row 33a is provided in the axial direction Da. is there.
- a plurality of stationary blades 47 constituting the fourth stage stationary blade row 46a which are low pressure components (second components), are arranged in the region where the blade ring air flow path 65f exists in the axial direction Da. Will be. Therefore, as in the above embodiments, the low-pressure component is lower than the pressure of the compressed air at the outlet of the intermediate compression stage 19a even if it is not arranged on the downstream side Dad in the axial direction from the blade ring air flow path. It only has to be arranged below.
- this embodiment is a modification of the first embodiment, in the second embodiment and the fourth to sixth embodiments as well, the downstream end in the axial direction of the blade ring air channel is the same as the present embodiment.
- the position of Ed may be shifted to the axial downstream side Dad.
- the gas turbine 1g of the present embodiment is a modification of the gas turbine 1f of the seventh embodiment.
- the low-pressure component (second component) in the seventh embodiment is only the stationary blade 47 that constitutes the fourth-stage stationary blade row 46a disposed on the downstream side Dad in the axial direction of the blade ring air passage 65f of the blade ring 44. .
- the low pressure component (second component) of the present embodiment includes a stationary blade 47 constituting the fourth stage stationary blade row 46a, a stationary blade 47 constituting the third stage stationary blade row 46b, and a radially outer Dro of the third stage stationary blade row 33b. It is the split ring 42 arrange
- a casing main body 45 that is a component of the turbine casing 41 of the present embodiment includes a cylindrical trunk portion 41a centering on an axis Ar, and a plurality of partition portions that extend axially inward from the inner peripheral surface of the trunk portion 41a. 41b.
- the some partition part 41b is located in a line with the axial direction Da at intervals in the axial direction Da.
- a blade ring 44 is attached to the end of each partition portion 41b on the radially inner side Dri.
- the space between the trunk portion 41a of the casing body 45 and the blade ring 44 in the radial direction Dr is partitioned into a plurality of spaces by a plurality of partition portions 41b.
- the present embodiment there are four partition portions 41b between the body portion 41a and the blade ring 44 in the radial direction Dr.
- three spaces are formed between the body portion 41a and the blade ring 44 in the radial direction by the four partition portions 41b.
- the first partition portion 41b1 on the most upstream side in the axial direction Dau and the second partition portion 41b2 adjacent to the downstream side in the axial direction Dad with respect to the first partition portion 41b1 are first.
- a space S1 is formed.
- a second space S2 is formed between the second partition portion 41b2 and the third partition portion 41b3 adjacent to the second partition portion 41b2 on the axially downstream side Dad.
- a third space S3 is formed between the third partition portion 41b3 and the fourth partition portion 41b4 on the most downstream side Dad in the axial direction.
- the reusable piping 67 that constitutes a part of the component introduction line 66 communicates with the third space S3.
- a fourth stage stationary blade row 46a exists at the position of the radially inner side Dri of the third space S3.
- the blade ring 44 penetrates from the surface defining the third space S3 in the outer circumferential surface of the blade ring 44 to the inner circumferential surface of the blade ring 44, and passes through the blade ring first penetration toward the fourth stage stationary blade row 46a.
- a hole 68a is formed.
- a partition portion through hole 68b penetrating in the axial direction Da is formed.
- the blade ring 44 has a blade ring second through-hole extending from the surface defining the second space S2 in the outer peripheral surface of the blade ring 44 to the inner peripheral surface of the blade ring 44 toward the third stage stationary blade row 46b. 68c is formed. Further, the blade ring 44 has a blade ring third through-hole extending from the surface defining the second space S2 in the outer peripheral surface of the blade ring 44 to the inner peripheral surface of the blade ring 44 toward the split ring 42. 68d is formed.
- the third space S ⁇ b> 3, the blade ring first through hole 68 a, the second space S ⁇ b> 2, the blade ring second through hole 68 c, and the blade ring third through hole 68 d are part of the flow path of the component introduction line 66.
- An air flow path 68g is configured.
- Part of the bleed air that has flowed into the third space S3 is introduced into the stationary blade 47 constituting the fourth stage stationary blade row 46a through the blade ring first through hole 68a, and cools the stationary blade 47.
- the remainder of the bleed air that has flowed into the third space S3 flows into the second space S2 through the partitioning portion through hole 68b of the third partitioning portion 44b3.
- Part of the bleed air that has flowed into the second space S2 is guided into the stationary blade 47 constituting the third stage stationary blade row 46b through the blade ring second through hole 68c, and cools the stationary blade 47.
- the remainder of the bleed air that has flowed into the second space S2 is guided into the split ring 42 via the blade ring third through hole 68d, and the split ring 42 is cooled.
- the low pressure component cooled by the bleed air flowing through the component introduction line 66 may be not only one type of component but also a plurality of types of components.
- the ring 42 is a low pressure part.
- one type or two types of components may be deleted from these components.
- the stationary blade 47 constituting the second-stage stationary blade row 46 and the split ring 42 arranged on the radially outer side Dro of the second-stage moving blade row 33 may be low-pressure components.
- it is necessary that these parts are arranged in a pressure environment lower than the pressure of the compressed air at the outlet of the intermediate compression stage 19a.
- present embodiment is a modification of the seventh embodiment, but also in the first to sixth embodiments, a plurality of types of components may be used as low-pressure components as in the present embodiment.
Abstract
Description
本願は、2015年11月26日に、日本国に出願された特願2015-231053号に基づき優先権を主張し、この内容をここに援用する。
複数の圧縮段を有し、各圧縮段で空気を順次圧縮する圧縮機と、前記圧縮機で圧縮された圧縮空気中で燃料を燃焼させて、燃焼ガスを生成する燃焼器と、前記燃焼ガスにより軸線を中心として回転するタービンロータと、前記タービンロータを覆う筒状のタービン車室と、前記タービン車室の内周側に固定されている複数の静翼列と、を有するタービンと、複数の前記圧縮段のうちの中間圧縮段から圧縮空気を抽気空気として抽気し、前記タービン車室の一部を構成する第一部品に前記抽気空気を導く抽気ラインと、前記タービンを構成する部品のうちで前記第一部品と異なる第二部品に、前記第一部品を通った前記抽気空気を導く部品導入ラインと、を備え、前記第一部品内には、前記軸線が延びる軸方向に延び、前記抽気ラインからの前記抽気空気が流れる第一空気流路が形成され、前記第二部品は、前記中間圧縮段の出口における前記圧縮空気の圧力より低い圧力環境下に配置されている低圧部品であり、前記低圧部品には、前記部品導入ラインからの前記抽気空気が流れる第二空気流路が形成されている。
前記第一態様のガスタービンにおいて、前記低圧部品は、前記第一部品内の前記第一空気流路よりも、前記燃焼ガスが流れる側である軸方向下流側に配置されている。
前記第一態様のガスタービンにおいて、前記低圧部品は、前記軸方向で、前記第一部品内の前記第一空気流路が存在している領域内に配置されている。
前記第一から第三態様のいずれかのガスタービンにおいて、前記抽気ラインは、前記第一空気流路の軸方向下流端に接続され、前記部品導入ラインは、前記第一空気流路の軸方向上流端に接続されている。
前記第一から第四態様のいずれかのガスタービンにおいて、前記低圧部品の前記第二空気流路は、前記部品導入ラインからの前記抽気空気を前記燃焼ガスが流れる燃焼ガス流路へ流出できるよう形成されている。
前記第一から第五態様のいずれかのガスタービンにおいて、複数の前記静翼列は、それぞれ、前記軸線に対する周方向に並んでいる複数の静翼を有し、前記低圧部品は、複数の前記静翼列のうち、前記第一空気流路よりも軸方向下流側に配置されている静翼列を構成する複数の静翼である。
前記第六態様のガスタービンにおいて、前記静翼は、前記軸線に対する径方向に延びて翼形を成す翼体と、前記翼体の径方向外側に設けられている外側シュラウドと、前記翼体の径方向内側に設けられている内側シュラウドと、を有し、前記低圧部品を構成する複数の前記静翼には、前記部品導入ラインからの抽気空気が前記外側シュラウドから流入し、前記翼体を経て前記内側シュラウドから流出する前記第二空気流路が形成されている。
前記第一から第五態様のいずれかのガスタービンにおいて、前記タービンロータは、前記軸線を中心として前記軸方向に延びるロータ軸と、前記軸方向に間隔をあけて並んでいる複数の動翼列と、を有し、前記低圧部品は、複数の前記動翼列のうち、前記中間圧縮段の出口における前記圧縮空気の圧力より低い圧力環境下に配置されている少なくとも一の動翼列を構成する複数の動翼であり、前記ロータ軸には、前記ロータ軸の軸方向の端から、前記低圧部品を構成する複数の前記動翼にまで延びる第三空気流路が形成され、前記部品導入ラインは、前記ロータ軸の前記第三空気流路を含む。
前記第一から第八態様のいずれかのガスタービンにおいて、前記タービンロータは、前記軸線を中心として前記軸方向に延びるロータ軸と、前記軸方向に間隔をあけて並んでいる複数の動翼列と、を有し、前記タービン車室は、前記動翼列と前記軸線に対する径方向で対向する複数の分割環と、複数の前記分割環及び複数の前記静翼列を径方向外側から支持する翼環と、前記翼環を径方向外側から支持する車室本体と、を有し、前記第一部品は、前記翼環である。
前記第九態様のガスタービンにおいて、前記翼環は、前記第一空気流路が形成されている領域内では、前記軸方向に関して一体形成品である。
前記第一から第十態様のいずれかのガスタービンにおいて、前記抽気ラインと前記部品導入ラインとに接続され、前記第一空気流路をバイパスさせて、前記抽気ラインに流入した前記抽気空気を前記部品導入ラインに導くバイパスラインと、前記抽気ラインに流入した前記抽気空気を前記第一空気流路に流入させる定常状態と、前記抽気ラインに流入した前記抽気空気を前記バイパスライン及び前記部品導入ラインを経て前記低圧部品に流入させるバイパス状態との間で切り替わる切替器と、を備える。
前記第十一態様のガスタービンにおいて、前記タービンの起動後、前記ガスタービンの出力、又は前記出力に相関性を有するパラメータである出力相関値が第一値になるまで、前記切替器に対してバイパス状態になるよう指令を出力し、前記出力相関値が前記第一値を超えると、前記切替器に対して前記定常状態になるよう指令を出力する冷却制御器を備える。
前記第一から第十態様のいずれかのガスタービンにおいて、前記第一部品を加熱する加熱手段と、前記タービンの起動過程で、遅くとも前記タービンロータが昇速開始される時から、前記ガスタービンの出力、又は前記出力に相関性を有するパラメータである出力相関値が第一値になるまでの間、前記加熱手段に前記第一部品を加熱させ、前記出力相関値が前記第一値を超えると、前記加熱手段による前記第一部品の加熱を中止させる加熱制御器と、を備える。
前記第十三態様のガスタービンにおいて、前記加熱手段は、前記圧縮機の起動停止動作とは独立して前記第一部品を加熱可能な手段であり、前記加熱制御器は、前記タービンの起動過程で、前記タービンロータが昇速開始される時よりも予め定められた時間前から、前記出力相関値が前記第一値になるまでの間、前記加熱手段に前記第一部品を加熱させる。
前記第十四態様のガスタービンにおいて、前記加熱手段は、前記抽気ライン又は前記第一部品に設けられて、熱を発する加熱器である。
前記第十三又は第十四態様の前記ガスタービンにおいて、前記加熱手段は、前記抽気ラインに接続され、前記中間圧縮段から抽気される前記抽気空気よりも温度が高い空気が流れる高温空気ラインと、切替器と、を有し、前記切替器は、前記高温空気ラインからの空気を前記抽気ラインを介して前記第一空気流路に流入させる加熱状態と、前記抽気ラインに流入した前記抽気空気を前記第一空気流路に流入させ、前記高温空気ラインからの空気を前記抽気ラインに流入させない定常状態との間で切り替わり、前記加熱制御器は、前記切替器の動作を制御する。
複数の圧縮段を有し、各圧縮段で空気を順次圧縮する圧縮機と、前記圧縮機で圧縮された圧縮空気中で燃料を燃焼させて、燃焼ガスを生成する燃焼器と、前記燃焼ガスにより軸線を中心として回転するタービンロータと、前記タービンロータを覆う筒状のタービン車室と、前記タービン車室の内周側に固定されている複数の静翼列と、を有するタービンと、を備えるガスタービンの部品温度調節方法において、複数の前記圧縮段のうちの中間圧縮段から圧縮空気を抽気空気として抽気して、前記タービン車室の一部を構成する第一部品中に前記抽気空気を流す第一空冷工程と、前記タービンを構成する部品のうちで前記第一部品と異なる第二部品中に、前記第一部品を通った抽気空気を流す第二空冷工程と、を実行し、前記第二部品は、前記中間圧縮段の出口における前記圧縮空気の圧力より低い圧力環境下に配置されている低圧部品である。
前記第十七態様のガスタービンの部品温度調節方法において、前記低圧部品は、前記第一部品中で前記抽気空気が流れる第一領域よりも、前記燃焼ガスが流れる側である軸方向下流側に配置されている。
前記第十七態様のガスタービンの部品温度調節方法において、前記低圧部品は、軸方向で、前記第一部品中で前記抽気空気が流れる第一領域が存在する範囲内に配置されている。
前記第十七から第十九態様のいずれかのガスタービンの部品温度調節方法において、前記第一空冷工程では、前記第一部品内で、軸方向上流側に前記抽気空気を流す。
前記第十七から第二十態様のいずれかのガスタービンの部品温度調節方法において、前記第二空冷工程では、前記低圧部品に流した前記抽気空気を前記燃焼ガスが流れる燃焼ガス流路に流出させる。
前記第十七から第二十一態様のいずれかのガスタービンの部品温度調節方法において、複数の前記静翼列は、それぞれ、前記軸線に対する周方向に並んでいる複数の静翼を有し、前記低圧部品は、複数の前記静翼列のうち、前記中間圧縮段の出口における前記圧縮空気の圧力より低い圧力環境下に配置されている少なくとも一の静翼列を構成する複数の静翼である。
前記第二十二態様のガスタービンの部品温度調節方法において、前記静翼は、前記軸線に対する径方向に延びて翼形を成す翼体と、前記翼体の径方向外側に設けられている外側シュラウドと、前記翼体の径方向内側に設けられている内側シュラウドと、を有し、前記第二空冷工程では、前記第一部品を通った抽気空気を、前記低圧部品を構成する前記静翼の前記外側シュラウドから前記静翼内に流入させ、前記静翼の前記翼体を経て、前記静翼の前記内側シュラウドから流出させる。
前記第十七から第二十一態様のいずれかのガスタービンの部品温度調節方法において、前記タービンロータは、前記軸線を中心として前記軸方向に延びるロータ軸と、前記軸方向に間隔をあけて並んでいる複数の動翼列と、を有し、前記低圧部品は、複数の前記動翼列のうち、前記中間圧縮段の出口における前記圧縮空気の圧力より低い圧力環境下に配置されている少なくとも一の動翼列を構成する複数の動翼である。
前記第十七から第二十四態様のいずれかのガスタービンの部品温度調節方法において、前記タービンロータは、前記軸線を中心として前記軸方向に延びるロータ軸と、前記軸方向に間隔をあけて並んでいる複数の動翼列と、を有し、前記タービン車室は、前記軸線に対する径方向外側に位置して複数の前記動翼列と径方向で対向する複数の分割環と、複数の分割環を径方向外側から支持する翼環と、前記翼環を径方向外側から支持する車室本体と、を有し、前記第一部品は、前記翼環である。
前記第十七から第二十五態様のいずれかのガスタービンの部品温度調節方法において、前記ガスタービンの出力、又は前記出力に相関性を有するパラメータである出力相関値が第一値を超えると、前記第一空冷工程及び前記第二空冷工程を実行する。
前記第二十六態様のガスタービンの部品温度調節方法において、前記タービンの起動後に前記出力相関値が前記第一値になるまで、前記中間圧縮段からの抽気空気を前記第一部品を経ずに前記低圧部品に流す第三空冷工程を実行する。
前記第二十六又は第二十七態様のガスタービンの部品温度調節方法において、前記第一空冷工程及び前記第二空冷工程の実行中に、前記出力相関値が前記第一値より小さい第二値未満になると、前記第一空冷工程及び前記第二空冷工程を中止し、前記中間圧縮段からの抽気空気を前記第一部品を経ずに前記低圧部品に流す第三空冷工程を実行する。
前記第二十七又は第二十八態様のガスタービンの部品温度調節方法において、前記第三空冷工程では、前記低圧部品を通った前記抽気空気を、前記燃焼ガスが流れる燃焼ガス流路に流出させる。
前記第二十六態様のガスタービンの部品温度調節方法において、前記タービンの起動過程で、遅くとも前記タービンロータが昇速開始される時から、前記出力相関値が前記第一値になるまでの間、前記第一部品を加熱し、前記出力相関値が前記第一値を超えると、前記第一部品の加熱を中止する加熱工程を実行する。
前記第三十態様のガスタービンの部品温度調節方法において、前記ガスタービンは、前記圧縮機の起動停止動作とは独立して前記第一部品を加熱可能な加熱手段を備え、前記加熱工程では、前記タービンの起動過程で、前記タービンロータが昇速開始される時よりも予め定められた時間前から、前記出力相関値が前記第一値になるまでの間、前記第一部品を前記加熱手段により加熱する。
前記第三十一態様のガスタービンの部品温度調節方法において、前記加熱手段は、熱を発する加熱器である。
本発明に係るガスタービンの第一実施形態について、図1~図5を参照して説明する。
各動翼列33は、いずれかの静翼列46の軸方向下流側Dadに配置されている。複数の動翼列33は、いずれも、周方向Dcに並んでいる複数の動翼34で構成されている。
本発明に係るガスタービンの第二実施形態について、図6及び図7を参照して説明する。
本発明に係るガスタービンの第三実施形態について、図8を参照して説明する。
本発明に係るガスタービンの第四実施形態について、図9~図11を参照して説明する。
本発明に係るガスタービンの第五実施形態について、図12を参照して説明する。
本発明に係るガスタービンの第六実施形態について、図13を参照して説明する。
本発明に係るガスタービンの第七実施形態について、図14を参照して説明する。
本発明に係るガスタービンの第八実施形態について、図15を参照して説明する。
2:ガスタービンロータ
3:ガスタービン車室
9:発電機
10:圧縮機
11:圧縮機ロータ
12:ロータ軸
13:動翼列
14:動翼
15:圧縮機車室
16:静翼列
17:静翼
19:圧縮段
19a:中間圧縮段
19b:高圧圧縮段
20:燃焼器
25:燃料ライン
26:燃料調節弁
30:タービン
31:タービンロータ
32:ロータ軸
32c:ディスクキャビティ
33:動翼列
33a:第四段動翼列
34:動翼(低圧部品)
34p:動翼空気流路(第二空気流路)
35:翼体
36:プラットフォーム
37:翼根
41:タービン車室
41a:胴部
41b:仕切部
41p:燃焼ガス流路
42:分割環
43:遮熱環
44:翼環(第一部品)
45:車室本体
46:静翼列
46a:第四段静翼列
46b:第三段静翼列
47:静翼(低圧部品)
47p:静翼空気流路(第二空気流路)
48:翼体
49o:外側シュラウド
49i:内側シュラウド
51:中間車室
52:排気室
61,61a:抽気ライン
65,65b,65f:翼環吸気流路(第一空気流路、第一領域)
66,66a,66e:部品導入ライン
67:使い回し配管
68,68g:車室内空気流路
69:ロータ空気流路
71:バイパスライン
75:切替器(三方弁)
80:加熱器
81:電気ヒータ
82:ヒータ駆動回路
84:第二抽気ライン(高温空気ライン)
85:切替器
100,100c,100d:制御器
102,102c,102d:部品温度制御部(冷却制御器及び/又は加熱制
御器)
Claims (32)
- 複数の圧縮段を有し、各圧縮段で空気を順次圧縮する圧縮機と、
前記圧縮機で圧縮された圧縮空気中で燃料を燃焼させて、燃焼ガスを生成する燃焼器と、
前記燃焼ガスにより軸線を中心として回転するタービンロータと、前記タービンロータを覆う筒状のタービン車室と、前記タービン車室の内周側に固定されている複数の静翼列と、を有するタービンと、
複数の前記圧縮段のうちの中間圧縮段から圧縮空気を抽気空気として抽気し、前記タービン車室の一部を構成する第一部品に前記抽気空気を導く抽気ラインと、
前記タービンを構成する部品のうちで前記第一部品と異なる第二部品に、前記第一部品を通った前記抽気空気を導く部品導入ラインと、
を備え、
前記第一部品内には、前記軸線が延びる軸方向に延び、前記抽気ラインからの前記抽気空気が流れる第一空気流路が形成され、
前記第二部品は、前記中間圧縮段の出口における前記圧縮空気の圧力より低い圧力環境下に配置されている低圧部品であり、
前記低圧部品には、前記部品導入ラインからの前記抽気空気が流れる第二空気流路が形成されている、
ガスタービン。 - 請求項1に記載のガスタービンにおいて、
前記低圧部品は、前記第一部品内の前記第一空気流路よりも、前記燃焼ガスが流れる側である軸方向下流側に配置されている、
ガスタービン。 - 請求項1に記載のガスタービンにおいて、
前記低圧部品は、前記軸方向で、前記第一部品内の前記第一空気流路が存在している領域内に配置されている、
ガスタービン。 - 請求項1から3のいずれか一項に記載のガスタービンにおいて、
前記抽気ラインは、前記第一空気流路の軸方向下流端に接続され、
前記部品導入ラインは、前記第一空気流路の軸方向上流端に接続されている、
ガスタービン。 - 請求項1から4のいずれか一項に記載のガスタービンにおいて、
前記低圧部品の前記第二空気流路は、前記部品導入ラインからの前記抽気空気を前記燃焼ガスが流れる燃焼ガス流路へ流出できるよう形成されている、
ガスタービン。 - 請求項1から5のいずれか一項に記載のガスタービンにおいて、
複数の前記静翼列は、それぞれ、前記軸線に対する周方向に並んでいる複数の静翼を有し、
前記低圧部品は、複数の前記静翼列のうち、前記中間圧縮段の出口における前記圧縮空気の圧力より低い圧力環境下に配置されている少なくとも一の静翼列を構成する複数の静翼である、
ガスタービン。 - 請求項6に記載のガスタービンにおいて、
前記静翼は、前記軸線に対する径方向に延びて翼形を成す翼体と、前記翼体の径方向外側に設けられている外側シュラウドと、前記翼体の径方向内側に設けられている内側シュラウドと、を有し、
前記低圧部品を構成する複数の前記静翼には、前記部品導入ラインからの抽気空気が前記外側シュラウドから流入し、前記翼体を経て前記内側シュラウドから流出する前記第二空気流路が形成されている、
ガスタービン。 - 請求項1から5のいずれか一項に記載のガスタービンにおいて、
前記タービンロータは、前記軸線を中心として前記軸方向に延びるロータ軸と、前記軸方向に間隔をあけて並んでいる複数の動翼列と、を有し、
前記低圧部品は、複数の前記動翼列のうち、前記中間圧縮段の出口における前記圧縮空気の圧力より低い圧力環境下に配置されている少なくとも一の動翼列を構成する複数の動翼であり、
前記ロータ軸には、前記ロータ軸の軸方向の端から、前記低圧部品を構成する複数の前記動翼にまで延びる第三空気流路が形成され、
前記部品導入ラインは、前記ロータ軸の前記第三空気流路を含む、
ガスタービン。 - 請求項1から8のいずれか一項に記載のガスタービンにおいて、
前記タービンロータは、前記軸線を中心として前記軸方向に延びるロータ軸と、前記軸方向に間隔をあけて並んでいる複数の動翼列と、を有し、
前記タービン車室は、前記動翼列と前記軸線に対する径方向で対向する複数の分割環と、複数の前記分割環及び複数の前記静翼列を径方向外側から支持する翼環と、前記翼環を径方向外側から支持する車室本体と、を有し、
前記第一部品は、前記翼環である、
ガスタービン。 - 請求項9に記載のガスタービンにおいて、
前記翼環は、前記第一空気流路が形成されている領域内では、前記軸方向に関して一体形成品である、
ガスタービン。 - 請求項1から10のいずれか一項に記載のガスタービンにおいて、
前記抽気ラインと前記部品導入ラインとに接続され、前記第一空気流路をバイパスさせて、前記抽気ラインに流入した前記抽気空気を前記部品導入ラインに導くバイパスラインと、
前記抽気ラインに流入した前記抽気空気を前記第一空気流路に流入させる定常状態と、前記抽気ラインに流入した前記抽気空気を前記バイパスライン及び前記部品導入ラインを経て前記低圧部品に流入させるバイパス状態との間で切り替わる切替器と、
を備える、
ガスタービン。 - 請求項11に記載のガスタービンにおいて、
前記タービンの起動後、前記ガスタービンの出力、又は前記出力に相関性を有するパラメータである出力相関値が第一値になるまで、前記切替器に対してバイパス状態になるよう指令を出力し、前記出力相関値が前記第一値を超えると、前記切替器に対して前記定常状態になるよう指令を出力する冷却制御器を備える、
ガスタービン。 - 請求項1から10のいずれか一項に記載のガスタービンにおいて、
前記第一部品を加熱する加熱手段と、
前記タービンの起動過程で、遅くとも前記タービンロータが昇速開始される時から、前記ガスタービンの出力、又は前記出力に相関性を有するパラメータである出力相関値が第一値になるまでの間、前記加熱手段に前記第一部品を加熱させ、前記出力相関値が前記第一値を超えると、前記加熱手段による前記第一部品の加熱を中止させる加熱制御器と、
を備えるガスタービン。 - 請求項13に記載のガスタービンにおいて、
前記加熱手段は、前記圧縮機の起動停止動作とは独立して前記第一部品を加熱可能な手段であり、
前記加熱制御器は、前記タービンの起動過程で、前記タービンロータが昇速開始される時よりも予め定められた時間前から、前記出力相関値が前記第一値になるまでの間、前記加熱手段に前記第一部品を加熱させる、
ガスタービン。 - 請求項14に記載のガスタービンにおいて、
前記加熱手段は、前記抽気ライン又は前記第一部品に設けられて、熱を発する加熱器である、
ガスタービン。 - 請求項13又は14に記載のガスタービンにおいて、
前記加熱手段は、前記抽気ラインに接続され、前記中間圧縮段から抽気される前記抽気空気よりも温度が高い空気が流れる高温空気ラインと、切替器と、を有し、
前記切替器は、前記高温空気ラインからの空気を前記抽気ラインを介して前記第一空気流路に流入させる加熱状態と、前記抽気ラインに流入した前記抽気空気を前記第一空気流路に流入させ、前記高温空気ラインからの空気を前記抽気ラインに流入させない定常状態との間で切り替わり、
前記加熱制御器は、前記切替器の動作を制御する、
ガスタービン。 - 複数の圧縮段を有し、各圧縮段で空気を順次圧縮する圧縮機と、
前記圧縮機で圧縮された圧縮空気中で燃料を燃焼させて、燃焼ガスを生成する燃焼器と、
前記燃焼ガスにより軸線を中心として回転するタービンロータと、前記タービンロータを覆う筒状のタービン車室と、前記タービン車室の内周側に固定されている複数の静翼列と、を有するタービンと、
を備えるガスタービンの部品温度調節方法において、
複数の前記圧縮段のうちの中間圧縮段から圧縮空気を抽気空気として抽気して、前記タービン車室の一部を構成する第一部品中に前記抽気空気を流す第一空冷工程と、
前記タービンを構成する部品のうちで前記第一部品と異なる第二部品中に、前記第一部品を通った抽気空気を流す第二空冷工程と、
を実行し、
前記第二部品は、前記中間圧縮段の出口における前記圧縮空気の圧力より低い圧力環境下に配置されている低圧部品である、
ガスタービンの部品温度調節方法。 - 請求項17に記載のガスタービンの部品温度調節方法において、
前記低圧部品は、前記第一部品中で前記抽気空気が流れる第一領域よりも、前記燃焼ガスが流れる側である軸方向下流側に配置されている、
ガスタービンの部品温度調節方法。 - 請求項17に記載のガスタービンの部品温度調節方法において、
前記低圧部品は、軸方向で、前記第一部品中で前記抽気空気が流れる第一領域が存在する範囲内に配置されている、
ガスタービンの部品温度調節方法。 - 請求項17から19のいずれか一項に記載のガスタービンの部品温度調節方法において、
前記第一空冷工程では、前記第一部品内で、軸方向上流側に前記抽気空気を流す、
ガスタービンの部品温度調節方法。 - 請求項17から20のいずれか一項に記載のガスタービンの部品温度調節方法において、
前記第二空冷工程では、前記低圧部品に流した前記抽気空気を前記燃焼ガスが流れる燃焼ガス流路に流出させる、
ガスタービンの部品温度調節方法。 - 請求項17から21のいずれか一項に記載のガスタービンの部品温度調節方法において、
複数の前記静翼列は、それぞれ、前記軸線に対する周方向に並んでいる複数の静翼を有し、
前記低圧部品は、複数の前記静翼列のうち、前記中間圧縮段の出口における前記圧縮空気の圧力より低い圧力環境下に配置されている少なくとも一の静翼列を構成する複数の静翼である、
ガスタービンの部品温度調節方法。 - 請求項22に記載のガスタービンの部品温度調節方法において、
前記静翼は、前記軸線に対する径方向に延びて翼形を成す翼体と、前記翼体の径方向外側に設けられている外側シュラウドと、前記翼体の径方向内側に設けられている内側シュラウドと、を有し、
前記第二空冷工程では、前記第一部品を通った抽気空気を、前記低圧部品を構成する前記静翼の前記外側シュラウドから前記静翼内に流入させ、前記静翼の前記翼体を経て、前記静翼の前記内側シュラウドから流出させる、
ガスタービンの部品温度調節方法。 - 請求項17から21のいずれか一項に記載のガスタービンの部品温度調節方法において、
前記タービンロータは、前記軸線を中心として軸方向に延びるロータ軸と、前記軸方向に間隔をあけて並んでいる複数の動翼列と、を有し、
前記低圧部品は、複数の前記動翼列のうち、前記中間圧縮段の出口における前記圧縮空気の圧力より低い圧力環境下に配置されている少なくとも一の動翼列を構成する複数の動翼である、
ガスタービンの部品温度調節方法。 - 請求項17から24のいずれか一項に記載のガスタービンの部品温度調節方法において、
前記タービンロータは、前記軸線を中心として軸方向に延びるロータ軸と、前記軸方向に間隔をあけて並んでいる複数の動翼列と、を有し、
前記タービン車室は、前記軸線に対する径方向外側に位置して複数の前記動翼列と径方向で対向する複数の分割環と、複数の分割環を径方向外側から支持する翼環と、前記翼環を径方向外側から支持する車室本体と、を有し、
前記第一部品は、前記翼環である、
ガスタービンの部品温度調節方法。 - 請求項17から25のいずれか一項に記載のガスタービンの部品温度調節方法において、
前記ガスタービンの出力、又は前記出力に相関性を有するパラメータである出力相関値が第一値を超えると、前記第一空冷工程及び前記第二空冷工程を実行する、
ガスタービンの部品温度調節方法。 - 請求項26に記載のガスタービンの部品温度調節方法において、
前記タービンの起動後に前記出力相関値が前記第一値になるまで、前記中間圧縮段からの抽気空気を前記第一部品を経ずに前記低圧部品に流す第三空冷工程を実行する、
ガスタービンの部品温度調節方法。 - 請求項26又は27に記載のガスタービンの部品温度調節方法において、
前記第一空冷工程及び前記第二空冷工程の実行中に、前記出力相関値が前記第一値より小さい第二値未満になると、前記第一空冷工程及び前記第二空冷工程を中止し、前記中間圧縮段からの抽気空気を前記第一部品を経ずに前記低圧部品に流す第三空冷工程を実行する、
ガスタービンの部品温度調節方法。 - 請求項27又は28に記載のガスタービンの部品温度調節方法において、
前記第三空冷工程では、前記低圧部品を通った前記抽気空気を、前記燃焼ガスが流れる燃焼ガス流路に流出させる、
ガスタービンの部品温度調節方法。 - 請求項26に記載のガスタービンの部品温度調節方法において、
前記タービンの起動過程で、遅くとも前記タービンロータが昇速開始される時から、前記出力相関値が前記第一値になるまでの間、前記第一部品を加熱し、前記出力相関値が前記第一値を超えると、前記第一部品の加熱を中止する加熱工程を実行する、
ガスタービンの部品温度調節方法。 - 請求項30に記載のガスタービンの部品温度調節方法において、
前記ガスタービンは、前記圧縮機の起動停止動作とは独立して前記第一部品を加熱可能な加熱手段を備え、
前記加熱工程では、前記タービンの起動過程で、前記タービンロータが昇速開始される時よりも予め定められた時間前から、前記出力相関値が前記第一値になるまでの間、前記第一部品を前記加熱手段により加熱する、
ガスタービンの部品温度調節方法。 - 請求項31に記載のガスタービンの部品温度調節方法において、
前記加熱手段は、熱を発する加熱器である、
ガスタービンの部品温度調節方法。
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Cited By (3)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
CN109083690A (zh) * | 2017-06-13 | 2018-12-25 | 通用电气公司 | 具有可变有效喉道的涡轮发动机 |
CN110030045A (zh) * | 2018-01-12 | 2019-07-19 | 通用电气公司 | 具有环形腔的涡轮发动机 |
WO2022054777A1 (ja) * | 2020-09-08 | 2022-03-17 | 三菱重工業株式会社 | ガスタービンのクリアランス制御システム |
Families Citing this family (7)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
JP6563312B2 (ja) * | 2015-11-05 | 2019-08-21 | 川崎重工業株式会社 | ガスタービンエンジンの抽気構造 |
US10920673B2 (en) * | 2017-03-16 | 2021-02-16 | General Electric Company | Gas turbine with extraction-air conditioner |
US11060463B2 (en) * | 2018-01-08 | 2021-07-13 | Raytheon Technologies Corporation | Modulated combustor bypass and combustor bypass valve |
FR3096071B1 (fr) | 2019-05-16 | 2022-08-26 | Safran Aircraft Engines | Contrôle de jeu entre des aubes de rotor d’aéronef et un carter |
JP7349320B2 (ja) * | 2019-10-25 | 2023-09-22 | 三菱重工業株式会社 | ガスタービン装置及びその製造方法並びにガスタービン装置の運転方法 |
WO2021194473A1 (en) * | 2020-03-24 | 2021-09-30 | Siemens Energy, Inc. | Method for modulating a turbine cooling supply for gas turbine applications |
US11668206B1 (en) * | 2022-03-09 | 2023-06-06 | General Electric Company | Temperature gradient control system for a compressor casing |
Citations (9)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
JPH1122413A (ja) * | 1997-07-08 | 1999-01-26 | Mitsubishi Heavy Ind Ltd | ガスタービン静翼のシール装置 |
JP2002213207A (ja) * | 2001-01-15 | 2002-07-31 | Mitsubishi Heavy Ind Ltd | ガスタービン分割環 |
US6487863B1 (en) * | 2001-03-30 | 2002-12-03 | Siemens Westinghouse Power Corporation | Method and apparatus for cooling high temperature components in a gas turbine |
WO2010084573A1 (ja) * | 2009-01-20 | 2010-07-29 | 三菱重工業株式会社 | ガスタービン設備 |
JP2012072708A (ja) * | 2010-09-29 | 2012-04-12 | Hitachi Ltd | ガスタービンおよびガスタービンの冷却方法 |
JP2012159056A (ja) * | 2011-02-02 | 2012-08-23 | Mitsubishi Heavy Ind Ltd | タービン静翼の流体供給構造 |
JP2012207565A (ja) * | 2011-03-29 | 2012-10-25 | Mitsubishi Heavy Ind Ltd | タービン排気構造及びガスタービン |
JP2013231439A (ja) * | 2012-05-01 | 2013-11-14 | General Electric Co <Ge> | 向流冷却システムを含むガスターボ機械及び方法 |
JP2014181702A (ja) * | 2013-03-15 | 2014-09-29 | General Electric Co <Ge> | ガスタービンのタービンセクションにおける熱伝達を改善するための方法および装置 |
Family Cites Families (23)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
JP3238299B2 (ja) | 1995-04-17 | 2001-12-10 | 三菱重工業株式会社 | ガスタービン |
US5611197A (en) * | 1995-10-23 | 1997-03-18 | General Electric Company | Closed-circuit air cooled turbine |
US5782076A (en) * | 1996-05-17 | 1998-07-21 | Westinghouse Electric Corporation | Closed loop air cooling system for combustion turbines |
WO1998058158A1 (fr) | 1997-06-19 | 1998-12-23 | Mitsubishi Heavy Industries, Ltd. | Dispositif d'etancheite pour aubes de stator de turbine a gaz |
JPH1193694A (ja) * | 1997-09-18 | 1999-04-06 | Toshiba Corp | ガスタービンプラント |
US6065282A (en) * | 1997-10-29 | 2000-05-23 | Mitsubishi Heavy Industries, Ltd. | System for cooling blades in a gas turbine |
WO2000060219A1 (de) | 1999-03-30 | 2000-10-12 | Siemens Aktiengesellschaft | Strömungsmaschine mit einer kühlbaren anordnung von wandelementen und verfahren zur kühlung einer anordnung von wandelementen |
DE10027833A1 (de) * | 2000-06-05 | 2001-12-13 | Alstom Power Nv | Verfahren zum Kühlen einer Gasturbinenanlage sowie Gasturbinenanlage zur Durchführung des Verfahrens |
US8495883B2 (en) * | 2007-04-05 | 2013-07-30 | Siemens Energy, Inc. | Cooling of turbine components using combustor shell air |
JP5101328B2 (ja) | 2008-02-12 | 2012-12-19 | 三菱重工業株式会社 | 軸流圧縮機およびこれを用いたガスタービン、ならびに抽気空気の冷却および熱回収方法 |
EP2674579B1 (en) * | 2008-10-08 | 2018-01-17 | MITSUBISHI HEAVY INDUSTRIES, Ltd. | Gas turbine and operating method thereof |
US8342798B2 (en) | 2009-07-28 | 2013-01-01 | General Electric Company | System and method for clearance control in a rotary machine |
WO2012001726A1 (ja) | 2010-06-28 | 2012-01-05 | 株式会社 日立製作所 | ガスタービンの間隙診断装置およびガスタービンシステム |
CH705181A1 (de) * | 2011-06-16 | 2012-12-31 | Alstom Technology Ltd | Verfahren zum Kühlen einer Gasturbinenanlage sowie Gasturbinenanlage zur Durchführung des Verfahrens. |
US9003807B2 (en) * | 2011-11-08 | 2015-04-14 | Siemens Aktiengesellschaft | Gas turbine engine with structure for directing compressed air on a blade ring |
US9222411B2 (en) * | 2011-12-21 | 2015-12-29 | General Electric Company | Bleed air and hot section component cooling air system and method |
US9541008B2 (en) * | 2012-02-06 | 2017-01-10 | General Electric Company | Method and apparatus to control part-load performance of a turbine |
JP5964106B2 (ja) | 2012-03-29 | 2016-08-03 | 三菱日立パワーシステムズ株式会社 | 圧縮機、及びガスタービン |
US9261022B2 (en) * | 2012-12-07 | 2016-02-16 | General Electric Company | System for controlling a cooling flow from a compressor section of a gas turbine |
US9714611B2 (en) * | 2013-02-15 | 2017-07-25 | Siemens Energy, Inc. | Heat shield manifold system for a midframe case of a gas turbine engine |
JP6189271B2 (ja) | 2013-09-20 | 2017-08-30 | 三菱重工業株式会社 | ガスタービン、ガスタービンの制御装置及びガスタービンの運転方法 |
JP6511734B2 (ja) | 2014-06-03 | 2019-05-15 | セイコーエプソン株式会社 | 原子セル、量子干渉装置、原子発振器、および電子機器 |
KR101790146B1 (ko) * | 2015-07-14 | 2017-10-25 | 두산중공업 주식회사 | 외부 케이싱으로 우회하는 냉각공기 공급유로가 마련된 냉각시스템을 포함하는 가스터빈. |
-
2016
- 2016-11-25 CN CN201680068684.6A patent/CN108291452B/zh active Active
- 2016-11-25 DE DE112016005433.5T patent/DE112016005433B4/de active Active
- 2016-11-25 US US15/778,122 patent/US10619564B2/en active Active
- 2016-11-25 WO PCT/JP2016/084916 patent/WO2017090709A1/ja active Application Filing
- 2016-11-25 JP JP2017552713A patent/JP6589211B2/ja active Active
- 2016-11-25 KR KR1020187014610A patent/KR101991645B1/ko active IP Right Grant
Patent Citations (9)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
JPH1122413A (ja) * | 1997-07-08 | 1999-01-26 | Mitsubishi Heavy Ind Ltd | ガスタービン静翼のシール装置 |
JP2002213207A (ja) * | 2001-01-15 | 2002-07-31 | Mitsubishi Heavy Ind Ltd | ガスタービン分割環 |
US6487863B1 (en) * | 2001-03-30 | 2002-12-03 | Siemens Westinghouse Power Corporation | Method and apparatus for cooling high temperature components in a gas turbine |
WO2010084573A1 (ja) * | 2009-01-20 | 2010-07-29 | 三菱重工業株式会社 | ガスタービン設備 |
JP2012072708A (ja) * | 2010-09-29 | 2012-04-12 | Hitachi Ltd | ガスタービンおよびガスタービンの冷却方法 |
JP2012159056A (ja) * | 2011-02-02 | 2012-08-23 | Mitsubishi Heavy Ind Ltd | タービン静翼の流体供給構造 |
JP2012207565A (ja) * | 2011-03-29 | 2012-10-25 | Mitsubishi Heavy Ind Ltd | タービン排気構造及びガスタービン |
JP2013231439A (ja) * | 2012-05-01 | 2013-11-14 | General Electric Co <Ge> | 向流冷却システムを含むガスターボ機械及び方法 |
JP2014181702A (ja) * | 2013-03-15 | 2014-09-29 | General Electric Co <Ge> | ガスタービンのタービンセクションにおける熱伝達を改善するための方法および装置 |
Cited By (6)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
CN109083690A (zh) * | 2017-06-13 | 2018-12-25 | 通用电气公司 | 具有可变有效喉道的涡轮发动机 |
CN110030045A (zh) * | 2018-01-12 | 2019-07-19 | 通用电气公司 | 具有环形腔的涡轮发动机 |
US11280198B2 (en) | 2018-01-12 | 2022-03-22 | General Electric Company | Turbine engine with annular cavity |
WO2022054777A1 (ja) * | 2020-09-08 | 2022-03-17 | 三菱重工業株式会社 | ガスタービンのクリアランス制御システム |
JP7339452B2 (ja) | 2020-09-08 | 2023-09-05 | 三菱重工業株式会社 | ガスタービンのクリアランス制御システム |
US11913341B2 (en) | 2020-09-08 | 2024-02-27 | Mitsubishi Heavy Industries, Ltd. | Clearance control system for gas turbine |
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