KR101617705B1 - Cooling System for Gas Turbine Vane and Blade Using Hybrid Coolant Cooling, Structure of Gas Turbine Vane and Blade, and Cooling Method Thereof - Google Patents
Cooling System for Gas Turbine Vane and Blade Using Hybrid Coolant Cooling, Structure of Gas Turbine Vane and Blade, and Cooling Method Thereof Download PDFInfo
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- KR101617705B1 KR101617705B1 KR1020150060320A KR20150060320A KR101617705B1 KR 101617705 B1 KR101617705 B1 KR 101617705B1 KR 1020150060320 A KR1020150060320 A KR 1020150060320A KR 20150060320 A KR20150060320 A KR 20150060320A KR 101617705 B1 KR101617705 B1 KR 101617705B1
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- cooling
- gas turbine
- fluid
- blade
- supercritical fluid
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- 238000001816 cooling Methods 0.000 title claims abstract description 120
- 239000002826 coolant Substances 0.000 title 1
- 239000007789 gas Substances 0.000 claims abstract description 128
- 239000012530 fluid Substances 0.000 claims abstract description 117
- 239000012809 cooling fluid Substances 0.000 claims abstract description 114
- 239000000567 combustion gas Substances 0.000 claims abstract description 5
- 239000000446 fuel Substances 0.000 claims abstract description 5
- 238000010248 power generation Methods 0.000 claims description 15
- 238000000034 method Methods 0.000 claims description 4
- 239000002131 composite material Substances 0.000 abstract description 4
- 230000008859 change Effects 0.000 abstract description 3
- CURLTUGMZLYLDI-UHFFFAOYSA-N Carbon dioxide Chemical compound O=C=O CURLTUGMZLYLDI-UHFFFAOYSA-N 0.000 description 30
- 229910002092 carbon dioxide Inorganic materials 0.000 description 15
- 239000001569 carbon dioxide Substances 0.000 description 15
- 230000006835 compression Effects 0.000 description 6
- 238000007906 compression Methods 0.000 description 6
- 238000009833 condensation Methods 0.000 description 5
- 230000005494 condensation Effects 0.000 description 5
- 238000010586 diagram Methods 0.000 description 5
- 238000011084 recovery Methods 0.000 description 5
- 238000002485 combustion reaction Methods 0.000 description 2
- 239000007788 liquid Substances 0.000 description 2
- 239000012528 membrane Substances 0.000 description 2
- 230000004048 modification Effects 0.000 description 2
- 238000012986 modification Methods 0.000 description 2
- 230000006866 deterioration Effects 0.000 description 1
- 230000007613 environmental effect Effects 0.000 description 1
- 230000006872 improvement Effects 0.000 description 1
- 230000035699 permeability Effects 0.000 description 1
- 230000008569 process Effects 0.000 description 1
- 230000009467 reduction Effects 0.000 description 1
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Classifications
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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
-
- 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
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
- F01D5/186—Film cooling
-
- 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
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
- F01D5/187—Convection cooling
-
- 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
- F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
- F01D9/023—Transition ducts between combustor cans and first stage of the turbine in gas-turbine engines; their cooling or sealings
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01K—STEAM ENGINE PLANTS; STEAM ACCUMULATORS; ENGINE PLANTS NOT OTHERWISE PROVIDED FOR; ENGINES USING SPECIAL WORKING FLUIDS OR CYCLES
- F01K25/00—Plants or engines characterised by use of special working fluids, not otherwise provided for; Plants operating in closed cycles and not otherwise provided for
- F01K25/08—Plants or engines characterised by use of special working fluids, not otherwise provided for; Plants operating in closed cycles and not otherwise provided for using special vapours
- F01K25/10—Plants or engines characterised by use of special working fluids, not otherwise provided for; Plants operating in closed cycles and not otherwise provided for using special vapours the vapours being cold, e.g. ammonia, carbon dioxide, ether
- F01K25/103—Carbon dioxide
-
- 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
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- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Chemical Kinetics & Catalysis (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
A gas turbine vane and blade cooling system using a combined cooling fluid is disclosed. A gas turbine vane and blade cooling system 100 using a composite cooling fluid according to an embodiment of the present invention includes a compressor 111 for compressing air, a combustor 112 for burning fuel and air exhausted from the compressor 111, And a turbine (113) driven by a combustion gas discharged from the combustor (112), wherein the compressed air and the supercritical fluid are flowed as a cooling fluid to cause the vanes (115) of the gas turbine A gas turbine system (100) for cooling a blade (116) of a gas turbine by means of film cooling and internal passage cooling, wherein compressed air is supplied as a cooling fluid And a supercritical fluid is used as a cooling fluid in a portion to which the internal flow path cooling is applied.
According to the gas turbine vane and blade cooling system of the present invention, by using the supercritical fluid having a density of the supercritical region and a supercritical fluid which is higher than the nonporous type compressed air and has no phase change in the supercritical region as a cooling fluid, can do.
Description
The present invention relates to a gas turbine vane and a blade cooling system using a composite cooling fluid, a gas turbine vane and a blade structure cooled thereby, and a cooling method for cooling a gas turbine vane and a blade using the same. More particularly, And a cooling system, a gas turbine vane, and a blade structure and a cooling method using the supercritical fluid as a cooling fluid.
As a power generation system according to the prior art, a gas turbine consists of a compressor, combustor turbine as a basic element. Specifically, the gas turbine compresses air with a compressor and directs the compressed air to the combustion chamber, where the fuel is injected and combusted. The high-temperature, high-pressure gas generated at this time is blown into the turbine while expanding to rotate the turbine.
Generally, the compressor and the turbine are directly or indirectly connected in one axis, and the power for operating the compressor uses 25 to 30% of the output generated by the turbine.
Therefore, the output for rotating the generator, the propeller, and the like with the gas turbine is obtained by subtracting the output required for operating the compressor from the output generated from the turbine.
FIG. 1 is a schematic diagram showing a method of cooling a combustor and a turbine in a gas turbine power generation system according to the prior art.
Referring to FIG. 1, a conventional gas turbine
However, since a part of the compressed cooling fluid is sent to the
It is an object of the present invention to provide a gas turbine vane and blade cooling system using a composite cooling fluid capable of improving the efficiency and cooling efficiency of a gas turbine system, a gas turbine vane and blade structure cooled thereby, And a cooling method for cooling the vane and the blade.
According to an aspect of the present invention, there is provided a gas turbine vane and blade cooling system comprising: a compressor for flowing compressed air and a supercritical fluid as a cooling fluid to perform a film cooling of a blade of a gas turbine and a blade of a gas turbine, And internal passage cooling, wherein compressed air is used as a cooling fluid in a portion to which the film cooling is applied, and a supercritical portion It may be to use the fluid as the cooling fluid.
According to an aspect of the present invention, there is provided a gas turbine vane and blade cooling system comprising: a compressor for flowing compressed air and a supercritical fluid as a cooling fluid to perform a film cooling of a blade of a gas turbine and a blade of a gas turbine, And cooling by internal passage cooling, wherein the gas turbine vane uses a supercritical fluid as a cooling fluid and the gas turbine blade uses compressed air as a cooling fluid Lt; / RTI >
In one embodiment of the present invention, the gas turbine system includes a supercritical fluid flowing through a cooling channel provided in a combustor and a turbine as a cooling fluid, a supercritical fluid flowing in the cooling channel, A supercritical fluid turbine system including a supercritical fluid compressor that compresses a supercritical fluid that has passed through a critical fluid turbine and a supercritical fluid turbine.
At this time, the supercritical fluid may be carbon dioxide.
In one embodiment of the present invention, an internal flow path may be formed in the vanes and the blades in an outward direction from the rotor shaft center so that the cooling fluid can flow.
In some cases, two or more internal passages may be formed in the vanes and the blades so that the cooling fluid can flow.
At this time, among the two or more internal flow paths, compressed air can be flowed as a cooling fluid in the internal flow path adjacent to the portion to which the film cooling is applied, and supercritical fluid can flow as the cooling fluid in the remaining internal flow paths.
In addition, a multi-layer wall with micro structure may be formed on the inner wall surface of the inner flow path which flows as supercritical fluid as a cooling fluid.
Further, the flow direction of the supercritical fluid may be a direction directed outward from the rotation center of the rotor shaft.
According to an aspect of the present invention, there is provided a gas turbine vane and a blade structure, wherein: a plurality of inner flow paths are formed inside the gas turbine vane and the blade, Among the two or more internal flow paths, compressed air may flow as a cooling fluid in the internal flow path adjacent to the film cooling portion, and supercritical fluid may flow as a cooling fluid in the remaining internal flow paths.
At this time, the portion to which the film cooling is applied may be a leading edge portion of the gas turbine vane and the blade.
According to another aspect of the present invention, there is provided a gas turbine vane and a blade structure, wherein: a plurality of internal flow paths are formed inside the gas turbine vane and the blade so as to allow cooling fluid to flow outwardly from a center of the rotor shaft , A supercritical fluid may be flowed as a cooling fluid in the internal flow path of the gas turbine vane, and compressed air may be flowed as a cooling fluid in the internal flow path of the gas turbine blade.
In an embodiment of the present invention, a multi-layer wall with micro structure may be formed on the inner wall surface of the inner flow path which flows as supercritical fluid as a cooling fluid.
In this case, the internal flow path may include: a supercritical fluid main flow path formed with a width W of a predetermined size; And a microstructure layer adjacent to the supercritical fluid main flow path, wherein the microstructure layer is formed by stacking two or more layers on the inner wall surface of the inner flow path and having a plurality of micro-passages formed therein;
. ≪ / RTI >
At this time, the thickness T formed by the microstructure layer may be 5 to 50% of the width W of the supercritical fluid main flow path.
A gas turbine vane and blade cooling method according to one aspect of the present invention includes the steps of: a) introducing a cooling fluid into a gas turbine vane and an internal flow passage formed in the blade, b) a cooling fluid recovery step of recovering the heated cooling fluid flowing through the gas turbine vane and an internal flow path formed inside the blade; c) driving the supercritical fluid turbine using the recovered cooling fluid; d) a supercritical fluid compression step of compressing the cooling fluid passing through the supercritical fluid turbine; And e) a cooling fluid circulating step of introducing the compressed cooling fluid into the gas turbine vane and an internal flow path formed inside the blade.
In one embodiment of the present invention, the cooling fluid inflow step includes the steps of: using compressed air as a cooling fluid in the gas turbine vane and a portion of the blade to which the membrane cooling is applied, wherein the gas turbine vane and the internal flow path cooling of the blade Supercritical fluid can be used as the cooling fluid in the part to which it is applied.
According to an embodiment of the present invention, the cooling fluid inflow step may include: using a supercritical fluid as a cooling fluid in an internal flow path of the gas turbine vane 115,
Compressed air can be used as a cooling fluid in the internal flow path of the gas turbine blade 116.
In one embodiment of the present invention, the supercritical fluid may be carbon dioxide.
As described above, according to the gas turbine vane and blade cooling system of the present invention, the compressed air and the supercritical fluid are flowed as the cooling fluid, so that the vanes of the gas turbine and the blades of the gas turbine are cooled by the film cooling, In a gas turbine system in which cooling is performed by internal passage cooling, high cooling efficiency is achieved by using compressed air as a cooling fluid in a film cooling portion and using a supercritical fluid as a cooling fluid in an inner flow path cooling portion .
Further, according to the gas turbine vane and the blade cooling system of the present invention, a supercritical fluid is used as a cooling fluid in a gas turbine vane having no influence of damage due to condensation due to no rotation, and a high risk of damage By using compressed air as the cooling fluid for the gas turbine blades, high cooling efficiency can be achieved.
Further, according to the gas turbine vane and blade cooling system of the present invention, since the supercritical fluid is used as the cooling fluid, the cooling performance is significantly improved as compared with the case where only the compressed air is used as the cooling fluid, The turbine vane portion may be substituted for the internal tube cooling method that includes a microfluid instead of the membrane cooling method.
Further, according to the gas turbine vane and blade cooling system of the present invention, by using the supercritical fluid having a higher density and specific heat of the supercritical region than the existing compressed air and no phase change in the supercritical region as the cooling fluid, Can be achieved.
Further, according to the gas turbine vane and blade cooling system of the present invention, a part of the cooling fluid necessary for the existing gas turbine cooling can be supplied from the supercritical fluid turbine system, thereby reducing the gas turbine system efficiency deterioration caused by the cooling fluid loss And as a result, the efficiency of the gas turbine system compared to the prior art can be improved.
Further, according to the gas turbine vane and blade structure of the present invention, the multi-layered micro-tube structure is formed on the inner wall surface of the partitioned and formed inner flow path, so that the gas turbine vane and the blade can be cooled more effectively, The efficiency can be improved.
In addition, according to the gas turbine vane and blade cooling method of the present invention, a cooling fluid inflow step, a cooling fluid recovery step, a supercritical fluid turbine driving step, a supercritical fluid compression step and a cooling fluid circulation step are provided, The cooling efficiency of the gas turbine system can be remarkably improved by using compressed air as the cooling fluid in the film cooling portion of the blade and using the supercritical fluid as the cooling fluid in the gas turbine vane and the inner flow path cooling portion of the blade.
BRIEF DESCRIPTION OF THE DRAWINGS FIG. 1 is a schematic diagram showing a method for cooling a combustor and a turbine in a gas turbine power generation system according to the prior art; FIG.
2 is a configuration diagram illustrating a gas turbine vane and blade cooling system in accordance with an embodiment of the present invention.
3 is a perspective view showing a structure of a gas turbine vane and a blade according to an embodiment of the present invention.
Fig. 4 is a cross-sectional view taken along the plane A of Fig. 3;
FIG. 5 is a view showing a flow direction of a cooling fluid flowing through an internal flow path formed inside the gas turbine vane and the blade shown in FIG. 3;
6 is an enlarged cross-sectional view of part B of Fig.
7 is a flow diagram illustrating a gas turbine vane and blade cooling method in accordance with an embodiment of the present invention.
Hereinafter, preferred embodiments of the present invention will be described in detail with reference to the drawings. Prior to the description, terms and words used in the present specification and claims should not be construed as limited to ordinary or dictionary meanings and should be construed in accordance with the technical concept of the present invention.
Throughout this specification, when a member is "on " another member, this includes not only when the member is in contact with another member, but also when there is another member between the two members.
Throughout this specification, when an element is referred to as "including" an element, it is understood that it may include other elements as well, without departing from the other elements unless specifically stated otherwise.
2 is a block diagram illustrating a gas turbine vane and blade cooling system in accordance with an embodiment of the present invention.
Referring to FIG. 2, the gas turbine vane and
Specifically, the main
In this case, compressed air may be used as a cooling fluid in a portion where film cooling is applied, and a supercritical fluid may be used as a cooling fluid in a portion where internal flow path cooling is applied.
In some cases, supercritical fluids are used as cooling fluids in gas turbine vanes, which are less susceptible to damage by condensation due to lack of rotation, and compressed air is used as cooling fluids in gas turbine blades, which have a high risk of damage due to rotation due to rotation Thereby achieving high cooling efficiency.
The supercritical
Generally, the compressed air used for the gas turbine vane and blade cooling fluid is about 400 ° C (the temperature for compressed air extracted from the compressor). Supercritical carbon dioxide can be applied to conventional gas turbine vanes and blade cooling fluids because the critical temperature of carbon dioxide is above 31 ° C and the operating temperature of the supercritical turbine can be varied depending on the supercritical fluid turbine system capacity. The gas turbine vane and blade cooling system according to the present embodiment is a system in which part of the cooling fluid required for existing gas turbine cooling is obtained from the supercritical system (15-25% of the total flow rate) The turbine system efficiency reduction can be reduced.
Supercritical fluid refers to a fluid at a point where it can not distinguish a liquid from a gas by reaching a state exceeding a certain high temperature and a high pressure limit. The density of the molecule is close to the liquid, but the viscosity is low.
At normal temperatures and pressures, when the temperature exceeds a certain high temperature and high pressure limit, which is referred to as a critical point, the process does not take place and thus the state can not be distinguished from the critical state. In other words, Fluid. Among the supercritical fluids having such properties, carbon dioxide is relatively close to room temperature. In the present invention, carbon dioxide is used as an embodiment of the supercritical fluid.
1, in the embodiment of the present invention, the cooling
In some cases, the
Therefore, according to the power generation system of the present invention, the carbon dioxide emitted from the combustion gas in the power generation system can be collected and recycled as a supercritical fluid, thereby preventing an environmental problem that may be generated due to the carbon dioxide exhaust gas.
FIG. 3 is a perspective view showing a structure of a gas turbine vane and a blade according to an embodiment of the present invention, and FIG. 4 is a sectional plane view taken along line A of FIG. 5 is a view showing a flow direction of a cooling fluid flowing through an internal flow path formed inside the gas turbine vane and the blade shown in Fig. 3, and Fig. 6 is an enlarged sectional view taken along line B of Fig. Respectively.
Referring to these drawings, an
In the two or more
In some cases, supercritical fluids are used as cooling fluids in gas turbine vanes, which are less susceptible to damage by condensation due to lack of rotation, and compressed air is used as cooling fluids in gas turbine blades, which have a high risk of damage due to rotation due to rotation Thereby achieving high cooling efficiency.
On the other hand, the flow direction of the supercritical fluid may be a direction from the rotation center of the rotor shaft to the outside as shown in Fig. In some cases, the flow direction of the supercritical fluid can be set differently according to the designer's intention.
6, a multi-layer wall with
More specifically, the
In some cases, the
Since the supercritical fluid operated in the supercritical
The thickness T of the
When the thickness T of the
7 is a flow chart illustrating a gas turbine vane and blade cooling method in accordance with an embodiment of the present invention.
Referring to FIG. 7 together with FIGS. 2 and 3, the cooling method S100 according to the present embodiment includes a cooling fluid inflow step S110, a cooling fluid recovery step S120, a supercritical fluid turbine driving step S130, , A supercritical fluid compression step (S140), and a cooling fluid circulation step (S150).
Specifically, the cooling fluid inflow step (S110) is a step of introducing the cooling fluid into the
At this time, compressed air is used as the cooling fluid in the portion where the gas cooling of the gas turbine vane 115 and the blade 116 is applied, and in the portion where the internal flow path cooling of the gas turbine vane 115 and the blade 116 is applied Supercritical fluid can be used as the cooling fluid. At this time, the above supercritical fluid may be carbon dioxide.
In some cases, supercritical fluids are used as cooling fluids in gas turbine vanes, which are less susceptible to damage by condensation due to lack of rotation, and compressed air is used as cooling fluids in gas turbine blades, which have a high risk of damage due to rotation due to rotation Thereby achieving high cooling efficiency.
In this case, since the supercritical fluid is used as the cooling fluid, the cooling performance is significantly improved as compared with the case where only the compressed air is used as the cooling fluid, and the gas turbine vane portion for flowing the supercritical fluid into the cooling fluid includes the micro- The internal tube cooling method can be applied instead.
The cooling fluid recovering step S120 is a step of recovering the heated cooling fluid flowing through the gas turbine vane 115 and the
The supercritical fluid turbine driving step (S130) is a step of driving the
The supercritical fluid compression step (S140) compresses the cooling fluid that has passed through the supercritical fluid turbine (121).
The cooling fluid circulation step (S150) is a step of introducing the compressed cooling fluid into an internal flow path formed inside the gas turbine vane 115 and the blade 116.
According to the gas turbine vane and blade cooling method (SlOO) of the present invention, the cooling fluid inflow step (S110), the cooling fluid recovery step (S120), the supercritical fluid turbine driving step (S130), the supercritical fluid compression step (S140) And a cooling fluid circulation step (S150), wherein compressed air is used as a cooling fluid in a portion where the gas cooling of the gas turbine vane and the blade is to be applied, and in a portion where internal flow cooling of the gas turbine vane and the blade is applied, By using the fluid as the cooling fluid, the cooling efficiency of the gas turbine system can be remarkably improved.
In the foregoing detailed description of the present invention, only specific embodiments thereof have been described. It is to be understood, however, that the invention is not to be limited to the specific forms thereof, which are to be considered as being limited to the specific embodiments, but on the contrary, the intention is to cover all modifications, equivalents, and alternatives falling within the spirit and scope of the invention as defined by the appended claims. .
That is, the present invention is not limited to the above-described specific embodiment and description, and various changes and modifications may be made without departing from the spirit and scope of the invention as defined in the appended claims. And such variations are within the scope of protection of the present invention.
100: Gas turbine vane and blade cooling system
110: main power generation system
111: Compressor
112: Combustor
113: Turbine
114:
115: Vane of gas turbine
116: Blade of gas turbine
117: leading edge
118: Trailing edge < RTI ID = 0.0 >
120: Supercritical fluid turbine system
121: Supercritical fluid turbine
122: supercritical fluid compressor
130:
131: a portion where film cooling is applied and an inner flow path
132: internal flow path to which internal flow path cooling is applied
133: multi-layer wall with micro structure
134: supercritical fluid main flow path
135: Micro through-hole
S100: Gas turbine vane and blade cooling method
S110: Cooling fluid inflow step
S120: cooling fluid recovery step
S130: Supercritical fluid turbine drive phase
S140: Supercritical fluid compression step
S150: cooling fluid circulation step
Claims (19)
Compressed air is used as a cooling fluid in the portion where the film cooling is applied,
Characterized in that a supercritical fluid is used as a cooling fluid in a portion to which the internal passage cooling is applied, and the cooling is performed by a film cooling and an internal passage cooling method by a gas turbine vane and blade cooling system A gas turbine vane and blade structure,
Inside the gas turbine vane 115 and the blades 116, a plurality of internal passages 130 are formed in the outer side from the center of the rotor shaft so that the cooling fluid can flow,
The compressed air is caused to flow as a cooling fluid in the inner flow path 131 adjacent to the film cooling portion of the two or more inner flow paths 130 and the supercritical fluid as a cooling fluid flows in the remaining inner flow path 132,
A multi-layer wall with microstructure 133 is formed on the inner wall surface of the inner flow path 132, which flows as supercritical fluid as a cooling fluid,
The internal flow path 132 includes:
A supercritical fluid main flow path 134 formed with a predetermined width W; And
A microstructure layer 133 formed by stacking two or more layers on the inner wall of the inner passage 132 adjacent to the supercritical fluid main flow passage 134 and having a plurality of micro passages 135 formed therein;
Wherein the gas turbine vane and blade structure comprise a plurality of gas turbine blades.
The gas turbine vane 115 uses a supercritical fluid as a cooling fluid,
Characterized in that the gas turbine blade (116) uses compressed air as a cooling fluid, and the gas turbine is cooled by a film cooling and an internal passage cooling method by a blade cooling system As the vane and blade structure,
Inside the gas turbine vane 115 and the blades 116, a plurality of internal passages 130 are formed in the outer side from the center of the rotor shaft so that the cooling fluid can flow,
A supercritical fluid is used as a cooling fluid in the internal flow path of the gas turbine vane 115,
In the internal flow path of the gas turbine blade 116, compressed air is used as a cooling fluid,
A multi-layer wall with microstructure 133 is formed on the inner wall surface of the inner flow path 132, which flows as supercritical fluid as a cooling fluid,
The internal flow path 132 includes:
A supercritical fluid main flow path 134 formed with a predetermined width W; And
And a microstructure layer 133 formed by stacking two or more layers on the inner wall surface of the inner channel 132 adjacent to the supercritical fluid main channel 134 and having a plurality of micro passageways 135 formed thereon Features a gas turbine vane and blade structure.
Wherein a thickness T formed by the microstructure layer 133 is 5 to 50% of a width W of the supercritical fluid main flow path 134. The gas turbine vane and blade structure according to claim 1,
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| KR1020150060320A KR101617705B1 (en) | 2015-04-29 | 2015-04-29 | Cooling System for Gas Turbine Vane and Blade Using Hybrid Coolant Cooling, Structure of Gas Turbine Vane and Blade, and Cooling Method Thereof |
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| KR1020150060320A KR101617705B1 (en) | 2015-04-29 | 2015-04-29 | Cooling System for Gas Turbine Vane and Blade Using Hybrid Coolant Cooling, Structure of Gas Turbine Vane and Blade, and Cooling Method Thereof |
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Cited By (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| KR101840324B1 (en) * | 2016-05-11 | 2018-03-21 | 한국기계연구원 | Gas turbin blade |
| WO2019054590A1 (en) * | 2017-09-12 | 2019-03-21 | 한국기계연구원 | Gas turbine blade |
| KR20190043740A (en) | 2017-10-19 | 2019-04-29 | 두산중공업 주식회사 | Gas turbine |
| WO2019245237A1 (en) | 2018-06-21 | 2019-12-26 | 한국기계연구원 | Gas turbine |
| KR20200145394A (en) | 2019-06-21 | 2020-12-30 | 두산중공업 주식회사 | Vane for turbine, turbine including the same |
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| JP2008261326A (en) * | 2007-04-10 | 2008-10-30 | General Electric Co <Ge> | Damper configured turbine blade |
| KR101485020B1 (en) * | 2013-12-12 | 2015-01-29 | 연세대학교 산학협력단 | Super critical fluid gas cooling gas turbine device |
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- 2015-04-29 KR KR1020150060320A patent/KR101617705B1/en active IP Right Grant
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| KR101840324B1 (en) * | 2016-05-11 | 2018-03-21 | 한국기계연구원 | Gas turbin blade |
| WO2019054590A1 (en) * | 2017-09-12 | 2019-03-21 | 한국기계연구원 | Gas turbine blade |
| US11313234B2 (en) | 2017-09-12 | 2022-04-26 | Korea Institute Of Machinery & Materials | Blade for gas turbine |
| KR20190043740A (en) | 2017-10-19 | 2019-04-29 | 두산중공업 주식회사 | Gas turbine |
| WO2019245237A1 (en) | 2018-06-21 | 2019-12-26 | 한국기계연구원 | Gas turbine |
| KR20190143625A (en) | 2018-06-21 | 2019-12-31 | 한국기계연구원 | Gas turbine |
| EP3812550A4 (en) * | 2018-06-21 | 2022-04-06 | Korea Institute of Machinery & Materials | Gas turbine |
| KR20200145394A (en) | 2019-06-21 | 2020-12-30 | 두산중공업 주식회사 | Vane for turbine, turbine including the same |
| US11299996B2 (en) | 2019-06-21 | 2022-04-12 | Doosan Heavy Industries & Construction Co., Ltd. | Turbine vane, and turbine and gas turbine including the same |
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