WO2018101190A1 - ガスタービン用高温部品、ガスタービンの翼及びガスタービン - Google Patents

ガスタービン用高温部品、ガスタービンの翼及びガスタービン Download PDF

Info

Publication number
WO2018101190A1
WO2018101190A1 PCT/JP2017/042332 JP2017042332W WO2018101190A1 WO 2018101190 A1 WO2018101190 A1 WO 2018101190A1 JP 2017042332 W JP2017042332 W JP 2017042332W WO 2018101190 A1 WO2018101190 A1 WO 2018101190A1
Authority
WO
WIPO (PCT)
Prior art keywords
porous
cooling gas
gas
gas turbine
distribution
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Ceased
Application number
PCT/JP2017/042332
Other languages
English (en)
French (fr)
Japanese (ja)
Inventor
石黒　達男
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Mitsubishi Heavy Industries Ltd
Original Assignee
Mitsubishi Heavy Industries Ltd
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Mitsubishi Heavy Industries Ltd filed Critical Mitsubishi Heavy Industries Ltd
Priority to US16/463,584 priority Critical patent/US20190277143A1/en
Publication of WO2018101190A1 publication Critical patent/WO2018101190A1/ja
Anticipated expiration legal-status Critical
Ceased legal-status Critical Current

Links

Images

Classifications

    • 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
    • F01D5/182—Transpiration cooling
    • F01D5/183—Blade walls being porous
    • 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
    • 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
    • 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
    • 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
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • 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
    • F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • 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
    • 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
    • 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
    • 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
    • F05D2240/20—Rotors
    • F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
    • 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
    • F05D2260/202—Heat transfer, e.g. cooling by film cooling
    • 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
    • F05D2300/00—Materials; Properties thereof
    • F05D2300/50—Intrinsic material properties or characteristics
    • F05D2300/514—Porosity

Definitions

  • the present disclosure relates to a high-temperature part for a gas turbine, a blade of a gas turbine, and a gas turbine.
  • the gas turbine includes a compressor, a combustor, and a turbine.
  • the gas turbine sucks in air, compresses the air with the compressor, burns fuel with the combustor, generates high-pressure and high-temperature combustion gas, and rotates the turbine. And electricity and thrust can be generated by the output of the turbine of the gas turbine.
  • High-temperature components for gas turbines such as a combustor, a turbine stationary blade, a moving blade, and a split ring are cooled by cooling air because they are exposed to high-temperature combustion gas.
  • a part applicable to the gas turbine disclosed in Patent Document 1 includes a substrate on which cooling air supply holes are formed, and a porous layer formed on the substrate. The porous layer is disposed on the gas path side through which the combustion gas flows, but is cooled by the cooling air supplied through the cooling air supply holes flowing inside the porous layer.
  • cooling air flows approximately in proportion to the square root of the difference between the inner pressure and the outer pressure (gas path static pressure).
  • gas path static pressure there is a pressure distribution on the outside (gas path side) of the high-temperature component for gas turbine.
  • Patent Document 1 if the distribution of the cooling air supply holes, the thickness of the porous layer and the porosity are constant, the flow rate of the cooling air is small in the region where the static pressure of the gas path is high, and the gas path The flow rate of the cooling air increases in a region where the static pressure is low. For this reason, an undesired distribution occurs in the flow rate of the cooling air in the porous layer.
  • the gas path static pressure on the ventral side is higher than the gas path static pressure on the back side, and the flow rate of cooling air on the ventral side is It becomes less and the abdominal surface side is more easily heated than the back surface side.
  • the component is a moving blade or a stationary blade, comparing the leading edge side and the trailing edge side, the gas path static pressure on the trailing edge side is lower than the gas path static pressure on the leading edge side, and the flow rate of the cooling air increases on the trailing edge side. Thereby, cooling air flows more than necessary from the rear edge side, and there is a possibility that the performance of the gas turbine is deteriorated.
  • the heat load acting on the high-temperature components for gas turbines there is a distribution in the heat load acting on the high-temperature components for gas turbines.
  • the heat load on the upstream side in the flow direction of the combustion gas is larger than that on the downstream side.
  • the distribution of the thermal load may also cause the component to be overheated locally.
  • it is not necessary to dispose a porous layer or cooling with cooling air may be unnecessary.
  • a cooling gas discharge passage is provided at the trailing edge of the blade of the gas turbine so that the cooling gas that has cooled the inside of the blade flows out from the trailing edge of the blade.
  • a cooling gas discharge channel it is necessary to increase the thickness of the trailing edge of the blade, which reduces the degree of freedom in designing the blade shape. there were.
  • the object of at least one embodiment of the present invention is to provide high-temperature components for gas turbines that prevent local overheating and gas turbine performance degradation due to excessive local flow of cooling gas, and It is to provide a gas turbine. Further, in view of the above-described circumstances, an object of at least one embodiment of the present invention is to allow the cooling gas that has cooled the inside of the blade to flow out from the trailing edge of the blade regardless of the thickness of the trailing edge of the blade. It is an object of the present invention to provide a gas turbine blade and a gas turbine.
  • a high temperature component for a gas turbine includes: The main body, A porous part provided as at least a part of the main body part or on at least a part of the main body part, through which a cooling gas can pass, According to one or both of the thermal load and pressure difference distribution acting on the main body part or the porous part, the distribution of the cooling gas in the porous part or the flow rate of the cooling gas in the porous part is distributed. It is configured.
  • the arrangement of the porous part or the flow rate of the cooling gas in the porous part Since it is configured to have a distribution, it is possible to prevent the main body portion and the porous portion from being overheated locally or the passage flow rate of the cooling gas from being excessively increased.
  • the porous part is provided on at least a part of the main body part,
  • the main body is provided with a plurality of cooling gas supply holes for supplying the cooling gas to the porous portion,
  • the distribution of the plurality of cooling gas supply holes is determined according to one or both of the thermal load and the pressure difference distribution acting on the porous portion.
  • the porous part is provided on at least a part of the main body part,
  • the main body is provided with a plurality of cooling gas supply holes for supplying the cooling gas to the porous portion,
  • the cross-sectional area of each of the plurality of cooling gas supply holes is determined in accordance with one or both of the thermal load and the pressure difference distribution acting on the porous portion.
  • the cross-sectional area of each of the plurality of cooling gas supply holes is determined according to one or both of the thermal load and the pressure difference distribution acting on the porous portion.
  • the main body is provided with a cavity having a cross-sectional area larger than that of the cooling gas supply hole between at least one of the plurality of cooling gas supply holes and the porous part.
  • the cooling gas is supplied to a wide area of the porous portion. Can be supplied. As a result, it is possible to prevent the porous portion from being overheated locally or the cooling gas passage flow rate from being excessively increased locally.
  • the porosity of the porous portion has a distribution according to one or both of the heat load and pressure difference distribution acting on the porous portion.
  • the porosity of the porous portion has a distribution according to one or both of the thermal load and the pressure difference distribution acting on the porous portion. It is prevented that the cooling gas is excessively heated and the flow rate of the cooling gas is excessively increased.
  • the thickness of the porous portion has a distribution according to one or both of the heat load and pressure difference distribution acting on the porous portion.
  • the thickness of the porous part has a distribution according to one or both of the thermal load and the pressure difference distribution acting on the porous part. It is prevented that the cooling gas is excessively heated and the flow rate of the cooling gas is excessively increased.
  • the main body portion or the porous portion constitutes at least a part of any one of a moving blade, a stationary blade, a split ring, and a combustor.
  • the porous portion is locally overheated or the cooling gas passage flow rate is locally excessive. Is prevented.
  • a blade of a gas turbine according to at least one embodiment of the present invention Consists of at least the trailing edge of the wing, and includes a porous part through which cooling gas can pass.
  • the porous part has a porosity distribution such that the cooling gas flows out from the rear edge of the wing part through the porous part from the inside of the wing part.
  • the trailing edge of the wing is configured by the porous portion and the cooling gas is allowed to flow through the porous portion. Cooling gas can be discharged from the trailing edge.
  • the cooling gas can be discharged from the trailing edge.
  • a gas turbine according to at least one embodiment of the present invention includes: The high temperature component for gas turbines as described in any one of the above configurations (1) to (7) is provided.
  • the porous portion in the gas turbine high-temperature component, the porous portion is prevented from being overheated locally and the passage flow rate of the cooling gas is prevented from becoming excessive locally. It can be operated at high temperatures. Thereby, a more efficient gas turbine is realizable.
  • a gas turbine according to at least one embodiment of the present invention includes: The blade of the gas turbine as described in said structure (8) is provided.
  • the cooling gas can be discharged from the trailing edge of the trailing edge, and the degree of freedom in designing the blades of the gas turbine is improved. As a result, it is possible to realize a gas turbine blade having a highly efficient shape, thereby realizing a highly efficient gas turbine.
  • a high-temperature component for a gas turbine and a gas turbine are provided in which local overheating and deterioration in performance of the gas turbine due to excessive local flow of cooling gas are prevented.
  • a gas turbine blade capable of allowing cooling gas cooled inside the blade to flow out of the blade trailing edge regardless of the thickness of the blade trailing edge, and A gas turbine is provided.
  • 1 is a diagram schematically illustrating a configuration of a gas turbine to which a high-temperature component for gas turbine according to an embodiment of the present invention is applied.
  • 1 is a perspective view schematically showing one stationary blade 11 applicable to a turbine as a high temperature component for a gas turbine according to an embodiment of the present invention.
  • 1 is a perspective view schematically showing one rotor blade applicable to a turbine as a high temperature component for a gas turbine according to an embodiment of the present invention. It is a perspective view showing roughly one division ring applicable to a turbine as a high temperature part for gas turbines concerning one embodiment of the present invention. It is a longitudinal cross-sectional view which shows schematically the split ring which concerns on one Embodiment of this invention.
  • FIG. 1 is a schematic cross-sectional view of a wing
  • (b) is a wing
  • 2 is a graph schematically showing the distribution of porosity in the porous portion on the ventral surface side and the distribution of porosity in the porous portion on the back surface side.
  • FIG. 1 is a cross-sectional view schematically showing a moving blade according to an embodiment of the present invention.
  • an expression indicating that things such as “identical”, “equal”, and “homogeneous” are in an equal state not only represents an exactly equal state, but also has a tolerance or a difference that can provide the same function. It also represents the existing state.
  • expressions representing shapes such as quadrangular shapes and cylindrical shapes represent not only geometrically strict shapes such as quadrangular shapes and cylindrical shapes, but also irregularities and chamfers as long as the same effects can be obtained. A shape including a part or the like is also expressed.
  • the expressions “comprising”, “comprising”, “comprising”, “including”, or “having” one constituent element are not exclusive expressions for excluding the existence of the other constituent elements.
  • FIG. 1 is a diagram schematically showing a configuration of a gas turbine 1 to which a high-temperature component for gas turbine according to an embodiment of the present invention is applied.
  • the gas turbine 1 includes a compressor (compression unit) 3, a combustor (combustion unit) 5, and a turbine (turbine unit) 7.
  • the compressor 3 sucks and compresses the atmosphere to generate compressed air.
  • the combustor 5 is supplied with compressed air from the compressor 3 together with the fuel, and the combustor 5 generates high-temperature and high-pressure combustion gas by burning the fuel.
  • the turbine 7 rotates the rotating shaft 9 using combustion gas.
  • the rotating shaft 9 is connected to the compressor 3 and is connected to, for example, a generator (not shown).
  • the compressor 3 is driven by the torque output from the rotating shaft 9 and the generator generates power.
  • FIG. 2 is a perspective view schematically showing one stationary blade 11 applicable to the turbine 7 as the high-temperature component 10 for a gas turbine according to the embodiment of the present invention.
  • the plurality of stationary blades 11 are fixed to a housing (cabinet) 12 of the turbine 7 in a state where the stationary blades 11 are arranged in the circumferential direction of the rotating shaft 9.
  • the stationary blade 11 includes a blade portion 13 and platforms 15 and 17 disposed on both sides of the blade portion 13, and a combustion gas flow path (gas path) is defined between the platforms 15 and 17. Therefore, the surfaces of the platforms 15 and 17 facing the gas path and the surface of the blade 13 are exposed to the combustion gas.
  • FIG. 3 is a perspective view schematically showing one rotor blade 19 applicable to the turbine 7 as the high-temperature component 10 for a gas turbine according to one embodiment of the present invention.
  • the plurality of rotor blades 19 are fixed to the rotating shaft 9 in a state of being arranged in the circumferential direction of the rotating shaft 9.
  • the moving blade 19 includes a blade portion 21, a platform 23 disposed on one side of the blade portion 21, and a blade root portion 25 that protrudes from the platform 23 to the opposite side of the blade portion 21. Since the blade root portion 25 is embedded in the rotating shaft 9, the moving blade 19 is fixed to the rotating shaft 9.
  • the platform 23 is disposed so as to cover the rotating shaft 9, and the surface of the platform 23 on the wing portion 21 side defines a gas path. Therefore, the surface of the platform 23 facing the gas path and the surface of the blade portion 21 are exposed to the combustion gas.
  • the combustion gas collides with the blade portions 21 of the plurality of moving blades 19 and rotates the rotating shaft 9.
  • FIG. 4 is a perspective view schematically showing one split ring 27 applicable to the turbine 7 as the high-temperature component 10 for a gas turbine according to the embodiment of the present invention.
  • the plurality of split rings 27 are fixed to the housing 12 of the turbine 7 while being arranged in the circumferential direction of the rotating shaft 9.
  • the split ring 27 is disposed outside the rotor blade 19 in the radial direction of the rotary shaft 9, and the plurality of split rings 27 arranged in the circumferential direction surround the plurality of rotor blades 19 arranged in the circumferential direction.
  • the split ring 27 includes a wall portion 29 that forms an surrounding wall that surrounds the rotor blade 19, and engaging portions 31 and 33 for fixing the wall portion 29 to the housing 12.
  • the surface (concave curved surface) of the wall 29 on the moving blade 19 side defines a gas path, and the surface of the wall 29 facing the gas path is exposed to the combustion gas.
  • FIG. 5 is a longitudinal sectional view schematically showing a split ring 27 (27a) according to an embodiment of the present invention.
  • FIG. 6 is a longitudinal sectional view schematically showing a split ring 27 (27b) according to an embodiment of the present invention.
  • FIG. 7 is a longitudinal sectional view schematically showing a split ring 27 (27c) according to an embodiment of the present invention.
  • FIG. 8 is a view for explaining a split ring 27 (27d) according to an embodiment of the present invention, (a) is a longitudinal sectional view schematically showing the split ring 27 (27d), and (b) ) Is a graph schematically showing the porosity distribution of the porous portion in the split ring 27 (27d).
  • FIG. 5 is a longitudinal sectional view schematically showing a split ring 27 (27a) according to an embodiment of the present invention.
  • FIG. 6 is a longitudinal sectional view schematically showing a split ring 27 (27b) according to an embodiment of the present invention.
  • FIG. 9 is a longitudinal sectional view schematically showing a split ring 27 (27e) according to an embodiment of the present invention.
  • FIG. 10 is a longitudinal sectional view schematically showing a split ring 27 (27f) according to an embodiment of the present invention.
  • FIG. 11 is a longitudinal sectional view schematically showing a split ring 27 (27g) according to an embodiment of the present invention.
  • FIG. 12 is a view for explaining a split ring 27 (27h) according to an embodiment of the present invention, (a) is a longitudinal sectional view schematically showing the split ring 27 (27h), and (b) ) Is a graph schematically showing the distribution of porosity in the split ring 27 (27h).
  • FIG. 13 is a longitudinal sectional view schematically showing a split ring 27 (27i) according to an embodiment of the present invention.
  • FIG. 14 is a longitudinal sectional view schematically showing a split ring 27 (27j) according to an embodiment of the present invention.
  • FIG. 15 is a view for explaining a moving blade 19 (19a) according to an embodiment of the present invention, (a) is a schematic cross-sectional view of the blade portion 21, and (b) is a combustion view.
  • 4 is a graph schematically showing a static pressure distribution on the outer side of the abdominal surface of the wing part 21, a static pressure distribution on the outer side of the back surface of the wing part 21, and a static pressure distribution on the inner side of the wing part 21 in the gas flow direction.
  • (C) is a graph which shows roughly the distribution of the porosity in the porous part of the abdominal surface side in the flow direction of combustion gas, and the distribution of the porosity in the porous part of the back side.
  • FIG. 16 is a view for explaining a moving blade 19 (19b) according to an embodiment of the present invention
  • (a) is a schematic cross-sectional view of the blade portion 21, and
  • (b) is a combustion view. It is a graph which shows roughly distribution of the heat load of the outer side of the abdominal surface of wing part 21, and the distribution of heat load of the back side outside of wing part 21 in the gas flow direction
  • (c) is a flow direction of combustion gas 2 is a graph schematically showing the distribution of porosity in the porous portion on the ventral side and the distribution of porosity in the porous portion on the back side.
  • FIG. 17 is a cross-sectional view schematically showing a rotor blade 19 (19c) according to an embodiment of the present invention.
  • FIG. 18 is a partial perspective view schematically showing a part of the blade portion 13 of the stationary blade 11 according to the embodiment of the present invention.
  • the high-temperature component for gas turbine 10 includes a main body portion 40 and a porous portion 42 as shown in FIGS.
  • the main body 40 constitutes a basic skeleton that forms the high-temperature component 10 for a gas turbine, and is constituted by, for example, a heat-resistant metal such as a Ni-based alloy, a ceramic matrix composite (CMC), or the like.
  • the CMC is formed, for example, a ceramic fiber such as SiC, Al 2 O 3, or the, by the ceramic matrix, such as, for example, SiC, Al 2 O 3, or the covering of ceramic fibers.
  • An intermediate layer such as BN is provided between the ceramic fiber and the ceramic matrix.
  • the porous part 42 is provided as at least a part of the main body part 40 or on at least a part of the main body part 40.
  • the porous part 42 constitutes a part of the wall of the high-temperature component 10 for gas turbine.
  • the porous part 42 constitutes a coating layer that covers the outer surface of the high-temperature component 10 for gas turbine.
  • the porous part 42 has minute pores (not shown), and the cooling gas can pass through the porous part 42 through the pores. That is, the porous part 42 has a fine cooling structure.
  • the cooling gas is, for example, air.
  • the porous part 42 is made of, for example, foam metal (porous metal) such as NiAl, ceramics such as porous yttrium-stabilized zirconia, or porous CMC.
  • the porous part 42 may be produced by a 3D printer, for example. Then, the gas turbine high-temperature component 10 is arranged in the porous portion 42 or cooled in the porous portion 42 according to one or both of the thermal load and the pressure difference distribution acting on the main body portion 40 or the porous portion 42.
  • the gas flow rate is configured to have a distribution.
  • the porous portion 42 is disposed on the gas path side in which the combustion gas flows in the high-temperature component 10 for gas turbine.
  • the high-temperature component 10 for gas turbine has an internal space 44 separated from a gas path by a porous portion 42, and a cooling gas having a pressure higher than that of the combustion gas flowing through the gas path is supplied to the internal space 44.
  • the porous portion 42 has an inner surface cooling or impingement cooling by cooling the main body portion 40 that defines the internal space 44, or a pressure difference between the static pressure of the cooling gas in the internal space 44 and the static pressure of the combustion gas in the gas path. Accordingly, the cooling gas is cooled by transfection cooling or microchannel cooling that passes through the porous portion 42.
  • the heat load and the pressure difference acting on the high-temperature component 10 for gas turbine do not act uniformly but have a distribution. Therefore, in the above configuration, the gas turbine high-temperature component 10 is arranged according to one or both of the thermal load acting on the main body 40 or the porous portion 42 and the distribution of the pressure difference, or the porous portion 42 is porous. Since the flow rate of the cooling gas in the portion 42 is configured to have a distribution, the main body 40 and the porous portion 42 may be locally overheated, or the flow rate of the cooling gas may be locally excessive. Is prevented.
  • the porous portion 42 is provided on at least a part of the main body 40. That is, the porous part 42 is formed in a layer shape so as to cover at least a part of the main body part 40 and constitutes a heat insulating film.
  • the main body 40 is provided with a plurality of cooling gas supply holes 46 for supplying cooling gas to the porous portion 42.
  • the cooling gas supply hole 46 fluidly connects the internal space 44 and the porous portion 42.
  • a plurality of cooling is performed according to one or both of the thermal load and the pressure difference distribution acting on the porous portion 42. The distribution of the gas supply holes 46 is determined.
  • the distribution of the cooling gas supply holes 46 is determined according to one or both of the thermal load and the pressure difference distribution acting on the porous portion 42. It is possible to prevent the mass portion 42 from being overheated locally and the passage flow rate of the cooling gas from becoming excessive excessively.
  • the heat load acting on the porous portion 42 becomes higher toward the upstream side in the flow direction of the combustion gas. Therefore, the cooling gas supply holes 46 are formed so that the number of the cooling gas supply holes 46 increases toward the upstream side in the flow direction of the combustion gas (so that the density of the cooling gas supply holes 46 increases). Further, for example, the pressure difference acting on the porous portion 42 becomes smaller toward the upstream side in the flow direction of the combustion gas. Therefore, the cooling gas supply holes 46 are formed so that the number of the cooling gas supply holes 46 increases toward the upstream side in the flow direction of the combustion gas. Further, for example, when the gas turbine high-temperature component 10 is a blade such as a stationary blade 11 or a moving blade 19, as shown in FIG.
  • the pressure difference on the ventral side is the smallest in the middle in the flow direction on the ventral side, and the largest in the middle in the flow direction on the back side. Therefore, according to such a pressure difference distribution, the number of cooling gas supply holes 46 increases as the pressure difference decreases on each of the back surface side and the abdominal surface side of the wing parts 13 and 21. The number of 46 may be determined.
  • the gas turbine high-temperature component 10 is a blade such as a stationary blade 11 or a moving blade 19, as shown in FIG. Distribution of heat load.
  • the heat load on the ventral side is highest at the leading edge in the flow direction, becomes minimal at the portion immediately downstream of the leading edge, gradually increases at the portion downstream, and further downstream. In this part, it gradually decreases to the trailing edge.
  • the heat load on the rear side is highest at the leading edge in the flow direction, becomes minimum at a portion immediately downstream of the leading edge, and gradually increases to the trailing edge at a portion downstream of the leading edge.
  • the cooling gas supply holes are arranged such that the number of the cooling gas supply holes 46 increases as the heat load increases on each of the rear surface side and the abdominal surface side of the wing parts 13 and 21.
  • the number of 46 may be determined.
  • the cooling gas supply hole 46 may be provided so as to supply the cooling gas only to the region of the porous portion 42 having a high heat load.
  • the cooling gas supply hole 46 may be provided only on the upstream side in the flow direction of the combustion gas.
  • the temperature of the main body portion 40 and the porous portion 42 can be kept below an allowable temperature only by cooling the high temperature component 10 for gas turbine from the inside. If possible, the cooling gas supply hole 46 may be omitted.
  • the porous portion 42 is provided on at least a part of the main body portion 40. That is, the porous part 42 is formed in a layer shape so as to cover at least a part of the main body part 40 and constitutes a heat insulating film.
  • the main body 40 is provided with a plurality of cooling gas supply holes 46 for supplying cooling gas to the porous portion 42.
  • the cooling gas supply hole 46 fluidly connects the internal space 44 and the porous portion 42.
  • each of the plurality of cooling gas supply holes 46 depends on one or both of the thermal load and the pressure difference distribution acting on the porous portion 42. In other words, the equivalent diameter is determined.
  • the cross-sectional area of each of the plurality of cooling gas supply holes 46 is determined according to one or both of the thermal load acting on the porous portion 42 and the distribution of the pressure difference.
  • the cooling gas supply hole 46 is formed so that the cross-sectional area of the cooling gas supply hole 46 increases toward the upstream side in the flow direction of the combustion gas.
  • the pressure difference acting on the porous portion 42 becomes smaller toward the upstream side in the flow direction of the combustion gas.
  • the cooling gas supply hole 46 is formed so that the cross-sectional area of the cooling gas supply hole 46 increases toward the upstream side in the flow direction of the combustion gas.
  • the gas turbine high-temperature component 10 is a blade such as a stationary blade 11 or a moving blade 19, as shown in FIG.
  • the pressure difference on the ventral side is the smallest in the middle in the flow direction on the ventral side, and the largest in the middle in the flow direction on the back side.
  • the cooling gas supply is performed so that the cross-sectional area of the cooling gas supply hole 46 increases as the pressure difference decreases on each of the back surface side and the abdominal surface side of the wing parts 13 and 21.
  • the cross-sectional area of the hole 46 may be determined.
  • the gas turbine high-temperature component 10 is a blade such as a stationary blade 11 or a moving blade 19, as shown in FIG. Distribution of heat load.
  • the heat load on the ventral side is highest at the leading edge in the flow direction, becomes minimal at the portion immediately downstream of the leading edge, gradually increases at the portion downstream, and further downstream. In this part, it gradually decreases to the trailing edge.
  • the heat load on the rear side is highest at the leading edge in the flow direction, becomes minimum at a portion immediately downstream of the leading edge, and gradually increases to the trailing edge at a portion downstream of the leading edge.
  • the cooling gas supply is performed so that the cross-sectional area of the cooling gas supply hole 46 increases as the heat load increases on each of the back surface side and the abdominal surface side of the wing portions 13 and 21.
  • the cross-sectional area of the hole 46 may be determined.
  • the main body 40 has a cooling gas supply hole 46 between at least one of the plurality of cooling gas supply holes 46 and the porous part 42.
  • a cavity 48 having a larger cross-sectional area is provided.
  • the cavity 48 is adjacent to the porous portion 42.
  • the cavity 48 has, for example, a cylindrical shape or a prism shape that is coaxial with the cooling gas supply hole 46.
  • the cavity 48 may have a groove shape or a channel shape extending along the porous portion 42.
  • the cavity 48 is formed so that the cross-sectional area of the cavity 48 is as large as possible.
  • the cavities 48 are formed such that the walls separating adjacent cavities 48 are as thin as possible.
  • the porous portion 42 is arranged according to one or both of the thermal load and the pressure difference distribution acting on the main body portion 40. According to the above configuration, since the porous portion 42 is arranged according to one or both of the thermal load and the pressure difference distribution acting on the main body portion 40, the porous portion 42 occupying the high-temperature component 10 for gas turbine. It is possible to protect the main body portion 40 from heat while reducing the ratio. For example, as shown in FIG. 10, the porous portion 42 is provided only on the upstream side in the flow direction of the combustion gas.
  • the porosity of the porous portion 42 is determined by the heat load and pressure acting on the porous portion 42. Depending on one or both of the difference distributions, it has a distribution.
  • the porosity of the porous portion 42 has a distribution according to one or both of the distribution of the thermal load and the pressure difference acting on the porous portion 42, so that the porous portion 42 is locally localized. It is prevented that the cooling gas is excessively heated and the flow rate of the cooling gas is excessively increased. For example, the heat load acting on the porous portion 42 becomes higher toward the upstream side in the flow direction of the combustion gas.
  • the porous portion 42 (42a to 42e) is formed so as to increase, for example, in a stepwise manner so that the porosity of the porous portion 42 increases toward the upstream side in the flow direction of the combustion gas. Further, for example, the pressure difference acting on the porous portion 42 becomes smaller toward the upstream side in the flow direction of the combustion gas. Therefore, the porous portion 42 (42a to 42e) is formed so as to increase, for example, in a stepwise manner so that the porosity of the porous portion 42 increases toward the upstream side in the flow direction of the combustion gas. Further, for example, when the gas turbine high-temperature component 10 is a blade such as a stationary blade 11 or a moving blade 19, as shown in FIG.
  • the pressure difference on the ventral side is the smallest in the middle in the flow direction on the ventral side, and the largest in the middle in the flow direction on the back side. Therefore, according to the distribution of the pressure difference, as shown in FIG. 15C, the pores of the porous part 42 become smaller as the pressure difference becomes smaller on the back side and the abdomen side of the wing parts 13 and 21.
  • the porosity distribution of the porous portion 42 may be determined so that the rate increases, for example, stepwise.
  • the gas turbine high-temperature component 10 is a blade such as a stationary blade 11 or a moving blade 19, as shown in FIG. Distribution of heat load.
  • the heat load on the ventral side is highest at the leading edge in the flow direction, becomes minimal at the portion immediately downstream of the leading edge, gradually increases at the portion downstream, and further downstream. In this part, it gradually decreases to the trailing edge.
  • the heat load on the rear side is highest at the leading edge in the flow direction, becomes minimum at a portion immediately downstream of the leading edge, and gradually increases to the trailing edge at a portion downstream of the leading edge. Therefore, according to the distribution of the heat load, as shown in FIG.
  • the pores of the porous portion 42 increase as the heat load increases on the back surface side and the abdominal surface side of the wing portions 13 and 21.
  • the porosity distribution of the porous portion 42 may be determined so that the rate increases, for example, in a stepwise manner.
  • the thickness of the porous portion 42 depends on one or both of the thermal load and pressure differential distribution acting on the porous portion 42, Have a distribution.
  • the porous portion 42 since the thickness of the porous portion 42 has a distribution according to one or both of the thermal load and the pressure difference distribution acting on the porous portion 42, the porous portion 42 is locally localized. It is prevented that the cooling gas is excessively heated and the flow rate of the cooling gas is excessively increased. For example, the heat load acting on the porous portion 42 becomes higher toward the upstream side in the flow direction of the combustion gas.
  • the porous portion 42 is formed so that the thickness of the porous portion 42 becomes thinner toward the upstream side in the flow direction of the combustion gas, for example, in a stepwise manner.
  • the pressure difference acting on the porous portion 42 becomes smaller toward the upstream side in the flow direction of the combustion gas. Therefore, the porous portion 42 is formed so that the thickness of the porous portion 42 becomes thinner toward the upstream side in the flow direction of the combustion gas, for example, in a stepwise manner.
  • the gas turbine high-temperature component 10 is a blade such as a stationary blade 11 or a moving blade 19, as shown in FIG. There is a distribution in the static pressure and a distribution in the pressure difference.
  • the pressure difference on the ventral side is the smallest in the middle in the flow direction on the ventral side, and the largest in the middle in the flow direction on the back side. Therefore, according to such a pressure difference distribution, the thickness of the porous portion 42 is reduced, for example, stepwise so that the pressure difference decreases on each of the back surface side and the abdominal surface side of the wing portions 13 and 21. You may determine distribution of the thickness of the porous part 42 so that it may become thin.
  • the gas turbine high-temperature component 10 is a blade such as a stationary blade 11 or a moving blade 19, as shown in FIG. Distribution of heat load.
  • the heat load on the ventral side is highest at the leading edge in the flow direction, becomes minimal at the portion immediately downstream of the leading edge, gradually increases at the portion downstream, and further downstream. In this part, it gradually decreases to the trailing edge.
  • the heat load on the rear side is highest at the leading edge in the flow direction, becomes minimum at a portion immediately downstream of the leading edge, and gradually increases to the trailing edge at a portion downstream of the leading edge.
  • the thickness of the porous portion 42 is reduced as the heat load increases on each of the back surface side and the abdominal surface side of the wing portions 13 and 21. You may determine distribution of the thickness of the porous part 42 so that it may become thin.
  • the porous portion 42 may constitute the entire main body portion 40 of the high-temperature component 10 for gas turbine.
  • a thermal barrier layer (TBC: Thermal Barrier Coating) 50 may be provided on the porous portion 42.
  • the heat shield layer 50 is made of, for example, ceramics such as yttrium stabilized zirconia, and has a porosity smaller than that of the porous portion 42.
  • a cooling gas discharge hole 52 through which the cooling gas flows out may be formed in the heat shield layer 50.
  • an adhesive layer (intermediate layer) 54 may be provided between the main body portion 40 and the porous portion 42 as shown in FIG.
  • the adhesive layer 54 joins the main body portion 40 and the porous portion 42, and is made of, for example, a material obtained by firing aluminum phosphate or an MCrAlY alloy.
  • M in the MCrAlY alloy represents one or more selected from the group consisting of Ni, Co, and Fe.
  • the MCrAlY alloy has a composition represented by Co-32Ni-21Cr-8Al-0.5Y.
  • the porous portion 42 is provided at least on the outer surface side (gas path side) of the wall portion 29 of the split ring 27. In some embodiments, as shown in FIGS. 15 to 17, the porous portion 42 is provided at least on the outer surface side (gas path side) of the blade portion 21 of the rotor blade 19. In some embodiments, as shown in FIG. 18, the porous portion 42 is provided at least on the outer surface side (gas path side) of the blade portion 13 of the stationary blade 11.
  • the main body portion 40 or the porous portion 42 constitutes at least a part of the moving blade 19, the stationary blade 11, or the split ring 27 as the high-temperature component 10 for a gas turbine.
  • the high temperature component 10 for gas turbines is the combustor 5 as shown in FIG.
  • the main body 40 or the porous portion 42 constitutes at least a part of the combustor 5, for example, a combustion cylinder or a tail cylinder (transition piece).
  • the main body portion 40 and the porous portion 42 are locally overheated or cooled in the same manner as the moving blade 19, the stationary blade 11 or the split ring 27. It is possible to prevent the gas flow rate from becoming excessive locally.
  • the blade of the gas turbine according to at least one embodiment of the present invention constitutes at least the rear edge portion 56 of the blade portion 21 as in the moving blade 19 shown in FIG.
  • Porous portions 42 and 58 through which gas can pass are provided.
  • the porous portions 42 and 58 have a porosity distribution such that the cooling gas flows out from the rear edge of the wing portion 21 through the porous portion 58 from the inside of the wing portion 21. More specifically, the porous portion (inner porous portion) 58 is covered with the porous portion (outer porous portion) 42, and the porosity of the porous portion 58 is higher than the porosity of the porous portion 42. large.
  • the cooling gas is easier to pass through the porous portion 58 than the porous portion 42, and the cooling gas is guided to the trailing edge of the wing portion 21 by the porous portion 58.
  • the thickness of the porous portion 42 is increased on the upstream side of the rear edge portion 56 so that the cooling gas does not flow out before the rear edge.
  • the trailing edge portion 56 of the wing portion 21 is configured by the porous portions 42 and 58, and the cooling gas is allowed to flow through the porous portion 58, so that the trailing edge portion 56 of the wing portion 21 is thin. Also, the cooling gas can be discharged from the trailing edge of the wing portion 21. At this time, by providing a distribution in the porosity of the porous portions 42 and 58 so that the cooling gas flows out from the rear edge of the wing portion 21, the rear edge portion before the cooling gas reaches the rear edge of the wing portion 21. It is possible to prevent all of the cooling gas from flowing out from the back side and the abdominal side of 56, and the cooling gas can be discharged from the rear edge.
  • the internal space 44 is partitioned into a plurality of compartments 60, and cooling gas is supplied to each compartment 60 in parallel.
  • the compartment 60 is entered.
  • a restrictor 62 is inserted in the cooling gas flow path. According to the above configuration, the static pressure of the cooling gas in the compartment 60 can be controlled by the throttle 62, and a large amount of cooling gas can be prevented from flowing out from the specific compartment 60.
  • FIG. 19 is a flowchart schematically showing an example of a procedure of a method for manufacturing a porous portion applied to the high-temperature component 10 for a gas turbine according to an embodiment of the present invention.
  • the manufacturing method of the porous portion includes a slurry preparation step S1, an assembly preparation step S3, a slurry applying step S5, a drying step S7, and a heating step S9.
  • a slurry preparation step S1 as a raw material of the slurry, water as a solvent, such as distilled water or deionized water, ceramic powder, pore-generating powder, a dispersant as necessary, and a binder as necessary And are prepared. Then, the raw materials are stirred and mixed to prepare a slurry.
  • the ceramic powder is, for example, a powder containing one or more selected from the group consisting of SiC, Si 3 N 4 , ⁇ SiAlON, AlN, TiB 2 , BN, WC, and the like, or a raw material thereof.
  • the pore-generating powder is, for example, a powder containing one or more selected from the group consisting of organic materials, carbon, graphite and the like.
  • the organic material powder is, for example, an acrylic, styrene, or polyethylene polymer powder.
  • Dispersants include, for example, polycarboxylic acid ammonium salt, polycarboxylic acid sodium salt, neutralized polyphosphate amino alcohol, naphthalenesulfonic acid ammonium salt, polycarboxylic acid alkylamine salt, nonionic surfactant, and cationic system 1 type or more selected from the group which consists of surfactant etc. is included.
  • the binder includes, for example, one or more selected from the group consisting of polyvinyl alcohol resin, acrylic resin, and paraffin.
  • the assembly preparation step S3 an assembly of ceramic fibers is prepared.
  • the aggregate of ceramic fibers is a bundle of ceramic fibers or a woven fabric.
  • the ceramic fiber includes, for example, one or more selected from the group consisting of SiC, SiTiCO, SiZrCO, SiAlCO, Si 3 N 4 and the like, or a raw material thereof.
  • slurry is applied to the ceramic fiber aggregate.
  • the slurry is applied to the aggregate of ceramic fibers so that the slurry penetrates into the gaps between the ceramic fibers.
  • the aggregate of ceramic fibers is immersed in the slurry under a pressure lower than atmospheric pressure.
  • the slurry is applied to the aggregate of ceramic fibers by applying the slurry and then rolling the roller.
  • the drying step S7 the slurry applied to the aggregate of ceramic fibers is dried in an atmosphere of, for example, 120 ° C. to form a green body (intermediate body).
  • the green body is heated in a reducing atmosphere of, for example, 1200 ° C., the ceramic powder is sintered, and the pore generating powder is lost.
  • the pore-generating powder is mixed in the slurry in the slurry preparation step S1, and the pore-generating powder is eliminated in the heating step S9. Pore corresponding to the powder can be generated.
  • a porosity can be controlled by adjusting the quantity of the powder for pore production
  • FIG. 20 is a schematic perspective view for explaining an example of the slurry applying step S5.
  • FIG. 21 is a schematic plan view for explaining an example of the slurry applying step S5.
  • slurry 68 is applied in parallel to ceramic fiber fabric 66 using a plurality of dispensers 64.
  • the plurality of dispensers 64 apply in parallel the slurry 68 having different contents of the pore-generating powder.
  • a roller 70 is applied along the extending direction of the slurry 68 to cause the slurry 68 to permeate the fabric 66.
  • slurry 68 having different contents of pore-generating powder is applied in parallel, and the roller 70 is applied along the extending direction of the slurry 68, so that the slurry 68 extends in a direction orthogonal to the extending direction of the slurry 68.
  • the roller 70 is applied along a direction orthogonal to the extending direction of the slurry 68, a porous part with a smooth change in porosity can be manufactured.
  • a plurality of fabrics 66 may be stacked depending on the thickness of the porous portion 42.
  • the operation of applying the slurry 68 to one fabric 66 by the dispenser 64, applying the roller 70, overlapping the next fabric 66, applying the slurry 68, and applying the roller 70 may be repeated.
  • the operation of applying the slurry 68 to one fabric 66 by the dispenser 64, stacking the next fabric 66, and applying the slurry 68 may be repeated, and finally the roller 70 may be applied collectively.
  • FIG. 22 is a schematic perspective view showing a plurality of fabrics 66 prepared in the assembly preparation step S3.
  • a plurality of fabrics 66 of different sizes are provided and stacked on top of each other. According to the above configuration, the thickness of the porous portion 42 can be changed with a simple configuration in accordance with the heat load and the pressure difference.
  • the split ring 27 has been mainly described as the high-temperature component 10 for a gas turbine.
  • the configuration described for the split ring 27 can be applied to the combustor 5, the stationary blade 11, and the moving blade 19.
  • the configuration described with reference to FIG. 5 can also be applied to the combustor 5, the rotor blade 19, and the split ring 27, and the configuration described with respect to the rotor blade 19 can be applied to the combustor 5, the stationary blade 11, and the split ring 27. .
  • the gas turbine high-temperature component 10 is a component that is at least partially heated to a temperature of, for example, 800 ° C. or more due to the influence of combustion gas.
  • the above-described combustor 5, stationary blade 11, moving blade 19, and The ring 27 is not limited.
  • the pattern of the slurry 68 applied to the fabric 66 in the slurry applying step S5 is not limited to being parallel as shown in FIGS. 20 and 21, and is appropriately selected according to the thermal load and the distribution of the pressure difference. Is possible.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
PCT/JP2017/042332 2016-11-30 2017-11-27 ガスタービン用高温部品、ガスタービンの翼及びガスタービン Ceased WO2018101190A1 (ja)

Priority Applications (1)

Application Number Priority Date Filing Date Title
US16/463,584 US20190277143A1 (en) 2016-11-30 2017-11-27 High-temperature component for gas turbine, gas turbine blade and gas turbine

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
JP2016-232044 2016-11-30
JP2016232044A JP6856364B2 (ja) 2016-11-30 2016-11-30 ガスタービン用高温部品、ガスタービンの翼及びガスタービン

Publications (1)

Publication Number Publication Date
WO2018101190A1 true WO2018101190A1 (ja) 2018-06-07

Family

ID=62242421

Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/JP2017/042332 Ceased WO2018101190A1 (ja) 2016-11-30 2017-11-27 ガスタービン用高温部品、ガスタービンの翼及びガスタービン

Country Status (3)

Country Link
US (1) US20190277143A1 (https=)
JP (1) JP6856364B2 (https=)
WO (1) WO2018101190A1 (https=)

Families Citing this family (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP6622176B2 (ja) * 2016-11-30 2019-12-18 三菱重工業株式会社 ガスタービン用高温部品及びガスタービン
EP3550106B1 (en) * 2018-04-06 2024-10-09 RTX Corporation Cooling air for gas turbine engine with thermally isolated cooling air delivery
US11913340B2 (en) * 2022-06-17 2024-02-27 Rtx Corporation Air seal system with backside abradable layer

Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP2010065634A (ja) * 2008-09-12 2010-03-25 Hitachi Ltd ガスタービンの高温部材
US7866948B1 (en) * 2006-08-16 2011-01-11 Florida Turbine Technologies, Inc. Turbine airfoil with near-wall impingement and vortex cooling
JP2011153540A (ja) * 2010-01-26 2011-08-11 Mitsubishi Heavy Ind Ltd 分割環冷却構造およびガスタービン
JP2013064365A (ja) * 2011-09-20 2013-04-11 Hitachi Ltd アブレイダブルコーティングを有する部材およびガスタービン
JP2014206154A (ja) * 2013-03-29 2014-10-30 ゼネラル・エレクトリック・カンパニイ タービン・システムのための熱ガス経路構成要素
JP2015048847A (ja) * 2013-08-30 2015-03-16 ゼネラル・エレクトリック・カンパニイ 多孔質冷却特徴部を有するガスタービン構成要素

Family Cites Families (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6511762B1 (en) * 2000-11-06 2003-01-28 General Electric Company Multi-layer thermal barrier coating with transpiration cooling

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US7866948B1 (en) * 2006-08-16 2011-01-11 Florida Turbine Technologies, Inc. Turbine airfoil with near-wall impingement and vortex cooling
JP2010065634A (ja) * 2008-09-12 2010-03-25 Hitachi Ltd ガスタービンの高温部材
JP2011153540A (ja) * 2010-01-26 2011-08-11 Mitsubishi Heavy Ind Ltd 分割環冷却構造およびガスタービン
JP2013064365A (ja) * 2011-09-20 2013-04-11 Hitachi Ltd アブレイダブルコーティングを有する部材およびガスタービン
JP2014206154A (ja) * 2013-03-29 2014-10-30 ゼネラル・エレクトリック・カンパニイ タービン・システムのための熱ガス経路構成要素
JP2015048847A (ja) * 2013-08-30 2015-03-16 ゼネラル・エレクトリック・カンパニイ 多孔質冷却特徴部を有するガスタービン構成要素

Also Published As

Publication number Publication date
JP2018087551A (ja) 2018-06-07
JP6856364B2 (ja) 2021-04-07
US20190277143A1 (en) 2019-09-12

Similar Documents

Publication Publication Date Title
EP2690260B1 (en) Seal segment
US6733907B2 (en) Hybrid ceramic material composed of insulating and structural ceramic layers
CN107250485B (zh) 具有保持热障涂层的专设表面特征的陶瓷基复合材料涡轮机部件
US10107119B2 (en) Vane assembly for a gas turbine engine
US20140294572A1 (en) Wall section for the working gas annulus of a gas turbine engine
US20060140753A1 (en) Blade outer seal with micro axial flow cooling system
US20180029944A1 (en) Ceramic matrix composite turbine component with engineered surface features retaining a thermal barrier coat
CN108138583B (zh) 燃气涡轮机燃烧器部件及其制造的方法和操作的方法
US20130260130A1 (en) Fiber-reinforced barrier coating, method of applying barrier coating to component and turbomachinery component
US20140127457A1 (en) Ceramic matrix composite component forming method
US11149948B2 (en) Fuel nozzle with angled main injection ports and radial main injection ports
EP3335873B1 (en) Gas turbine engine vane
JP6622176B2 (ja) ガスタービン用高温部品及びガスタービン
US11352886B2 (en) Coated components having adaptive cooling openings and methods of making the same
WO2018101190A1 (ja) ガスタービン用高温部品、ガスタービンの翼及びガスタービン
JP6067869B2 (ja) タービンエアロフォイルのアブレイダブル皮膜システムおよび対応するタービンブレード
US10294807B2 (en) Inter-turbine ducts
US20160333784A1 (en) A wall cooling arrangement for a gas turbine engine
US12601275B1 (en) CMC component with cooling cavity
US11286783B2 (en) Airfoil with CMC liner and multi-piece monolithic ceramic shell
US12467629B1 (en) CMC component with cover plate
WO2020023021A1 (en) Ceramic matrix composite materials having reduced anisotropic sintering shrinkage
WO2022035460A1 (en) Thermal barrier coating, layered component comprising a thermal barrier coating and gas turbine engine comprising a layered component
US11085640B2 (en) Production of a combustion chamber made of composite material

Legal Events

Date Code Title Description
121 Ep: the epo has been informed by wipo that ep was designated in this application

Ref document number: 17876972

Country of ref document: EP

Kind code of ref document: A1

NENP Non-entry into the national phase

Ref country code: DE

122 Ep: pct application non-entry in european phase

Ref document number: 17876972

Country of ref document: EP

Kind code of ref document: A1