US20200400026A1 - Ceramic matrix composite vane with trailing edge radial cooling - Google Patents
Ceramic matrix composite vane with trailing edge radial cooling Download PDFInfo
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- US20200400026A1 US20200400026A1 US16/448,137 US201916448137A US2020400026A1 US 20200400026 A1 US20200400026 A1 US 20200400026A1 US 201916448137 A US201916448137 A US 201916448137A US 2020400026 A1 US2020400026 A1 US 2020400026A1
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- aerofoil
- trailing
- matrix composite
- ceramic matrix
- edge
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Classifications
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—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
- F01D5/185—Liquid 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
- F01D5/147—Construction, i.e. structural features, e.g. of weight-saving hollow 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
- 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
- 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
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/28—Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
- F01D5/282—Selecting composite materials, e.g. blades with reinforcing filaments
-
- 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
- F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2220/00—Application
- F05D2220/30—Application in turbines
- F05D2220/32—Application in turbines in gas turbines
-
- 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
- F05D2240/12—Fluid guiding means, e.g. vanes
- F05D2240/122—Fluid guiding means, e.g. vanes related to the trailing edge of a stator vane
-
- 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
- F05D2260/221—Improvement of heat transfer
-
- 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
- F05D2300/60—Properties or characteristics given to material by treatment or manufacturing
- F05D2300/603—Composites; e.g. fibre-reinforced
- F05D2300/6033—Ceramic matrix composites [CMC]
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
Definitions
- the present disclosure relates generally to components for gas turbine engines, and more specifically to aerofoils that comprise ceramic-containing materials.
- the present disclosure may comprise one or more of the following features and combinations thereof.
- a component for a gas turbine engine includes a ceramic matrix composite aerofoil.
- the ceramic matrix composite aerofoil is adapted to conduct gases flowing through a gas path of a gas turbine engine around the component during use of the component.
- the ceramic matrix composite aerofoil includes a leading edge, a trailing edge spaced apart axially from the leading edge relative to an axis, a pressure side that extends between and interconnects the leading edge and the trailing edge, and a suction side that extends between and interconnects the leading edge and the trailing edge.
- the ceramic matrix composite aerofoil may be formed to define an aerofoil-shaped passage that extends radially at least partway into the ceramic matrix composite aerofoil.
- the trailing edge, an aft portion of the pressure side, and an aft portion of the suction side of the ceramic matrix composite aerofoil may define a trailing-edge passage.
- the trailing-edge passage may extend radially through the ceramic matrix composite aerofoil to conduct cooling fluid radially through the trailing edge of the ceramic matrix composite aerofoil.
- a minimum radius of the trailing-edge passage may be smaller than a minimum radius of an aft end of the aerofoil-shaped passage.
- the trailing edge, the aft portion of the pressure side, and the aft portion of the suction side of the ceramic matrix composite aerofoil may be formed without passages that provide fluid communication between the gas path and the trailing-edge passage.
- the trailing-edge passage may extend radially through the entire component.
- the ceramic matrix composite aerofoil may include an inner surface that forms the trailing-edge passage and the inner surface may be continuous and without holes.
- the trailing-edge passage and the aerofoil-shaped passage are the only voids in the ceramic matrix composite aerofoil.
- the aerofoil-shaped passage extends radially entirely through the ceramic matrix composite outer platform and the inner platform.
- the component may include a metallic support strut located in the aero-foil shaped passage of the ceramic matrix composite aerofoil and configured to receive force loads acting on the ceramic matrix composite aerofoil during use of the component.
- the ceramic matrix composite aerofoil includes a core body, a trailing edge filler, and an outer layer.
- the core body may define the leading edge, a fore portion of the pressure side, and a fore portion of the suction side of the ceramic matrix composite aerofoil.
- the trailing edge filler may define the trailing edge, the aft portion of the pressure side, and the aft portion of the suction side of the ceramic matrix composite aerofoil.
- the outer layer may extend around the core body and the trailing edge filler to provide an outermost surface of the ceramic matrix composite aerofoil.
- the trailing edge filler may be made of ceramic matrix composite materials.
- the trailing edge fill may be formed to define the entire trailing-edge passage.
- the core body and the outer layer may be made of ceramic matrix composite materials.
- the core body may define the entire aerofoil-shaped passage.
- the core body, the trailing edge filler, and the outer layer may be integrally formed such that the ceramic matrix composite aerofoil is a single, one-piece member.
- the ceramic matrix composite aerofoil may be formed to define a plurality of trailing-edge passages that includes the trailing-edge passage.
- the plurality of trailing-edge passages may include at least a circular shaped trailing-edge passage and an ellipse shaped trailing-edge passage.
- the trailing-edge passage may be ellipse shaped. In some embodiments, the trailing-edge passage may be generally triangular.
- a method may include a number of steps.
- the steps may include providing an aerofoil-shaped core preform that is formed to define an aerofoil-shaped passage that extends radially there through, providing a trailing edge filler preform that is formed to define a trailing-edge passage that extends radially there through, infiltrating the aerofoil-shaped core preform and the trailing edge filler preform together to provide a single, unitary ceramic matrix composite aerofoil component having the aerofoil-shaped passage and the trailing-edge passage, and directing a first portion of a cooling fluid radially through the aerofoil-shaped passage and directing a second portion of the cooling fluid radially through the trailing-edge passage.
- the aerofoil-shaped passage may have a minimum radius.
- the trailing-edge passage may have a minimum radius.
- the minimum radius of the trailing-edge passage may be smaller than the minimum radius of the aerofoil-shaped passage.
- the method may include directing the second portion of the cooling fluid from an outer chamber located radially outward of the ceramic matrix composite aerofoil component, radially inward through the trailing-edge passage, and into an inner chamber located radially inward of the ceramic matrix composite aerofoil.
- the method may include directing the second portion of the cooling fluid from an inner chamber located radially inward of the ceramic matrix composite aerofoil component, radially outward through the trailing-edge passage, and into an outer chamber located radially outward of the ceramic matrix composite aerofoil.
- the method may include machining an inner surface of the ceramic matrix composite aerofoil component that defines the trailing-edge passage to cause the inner surface to have a surface roughness that provides a desired heat transfer between the second portion of the cooling fluid and the inner surface.
- FIG. 1 is a cutaway view of a gas turbine engine that includes a fan, a compressor, a combustor, and a turbine, the turbine including static turbine vane rings configured to direct air into adjacent rotating wheel assemblies;
- FIG. 2 is a perspective view of a component of the turbine shown in FIG. 1 suggesting that the component includes a ceramic matrix composite aerofoil formed to include an aerofoil-shaped passage and a trailing-edge passage that extends radially through the trailing edge of the ceramic matrix composite aerofoil;
- FIG. 3 is a section view taken along line 3 - 3 of the component of FIG. 2 showing that the component further includes a support strut located in the aerofoil-shaped passage and the ceramic matrix composite aerofoil is made of an aerofoil core, a trailing edge filler formed to define the trailing-edge passage, and an outer layer arranged around the aerofoil core and the trailing edge filler;
- FIG. 4 is an enlarged view of the trailing edge of the component of FIG. 3 showing an aft end of the aerofoil-shaped passage and the trailing-edge passage formed in the component;
- FIG. 5 is a section view of the turbine of the gas turbine engine of FIG. 1 showing the component in the turbine;
- FIG. 6 is a section view of a trailing edge of another embodiment of a ceramic matrix composite aerofoil in accordance with the present disclosure showing that the trailing-edge passage is generally triangular;
- FIG. 7 is a section view of a trailing edge of another embodiment of a ceramic matrix composite aerofoil in accordance with the present disclosure showing that the trailing edge is formed to define a plurality of trailing-edge passages therein.
- a component 10 for use in a gas turbine engine 110 is shown in FIG. 2 .
- the illustrative component 10 includes a ceramic matrix composite aerofoil 12 as shown in FIGS. 2-5 .
- the ceramic matrix composite aerofoil 12 interacts with hot gases conducted through a gas path 15 of the gas turbine engine 110 and conducts the hot gases around the component 10 toward a rotating wheel assembly 22 located downstream of the component 10 .
- the component 10 is a blade included in the rotating wheel assembly 22 , a seal segment, or other suitable component.
- the aerofoil 12 is formed to include a passage 24 (sometimes called an internal cavity) and a trailing-edge passage 26 as shown in FIGS. 2 and 3 .
- the passage 24 extends radially at least partway into the aerofoil 12 and, illustratively, extends radially through the aerofoil 12 .
- the trailing-edge passage 26 is located at a trailing edge 30 of the aerofoil 12 and extends radially through the aerofoil 12 to provide cooling to the trailing edge 30 .
- some gas turbine engines have aerofoils with trailing edges that are cooled by axially extending cooling holes formed in the aerofoils and the cooling air is expelled to the gas path.
- the trailing-edge passage 26 of the present disclosure extends radially through the aerofoil 12 to reduce or remove the design constraints of the aerofoil 12 such as, for example, film air ejection (pressure, Mach no. location, rate) at the trailing edge 30 , trailing edge thickness, trailing edge length, and size and position of the passage 24 .
- the trailing-edge passage 26 is formed in a preform separate from the preform that defines the passage 24 .
- the trailing-edge passage 26 is not fluidly connected directly with the passage 24 via other holes or passages.
- the trailing-edge passage 26 is not fluidly connected with any axially extending holes that are open to the gas path 15 .
- the component 10 is a vane 40 in the illustrative embodiment and includes a metallic support strut 14 , a ceramic matrix composite outer platform 16 , and a ceramic matrix composite inner platform 18 as shown in FIG. 2 .
- the metallic support strut 14 extends through the passage 24 of the ceramic matrix composite aerofoil 12 and receives force loads applied to the ceramic matrix composite aerofoil 12 by the hot gases.
- the outer platform 16 defines an outer boundary of the gas path 15 and the inner platform is spaced apart radially from the outer platform 16 and defines an inner boundary of the gas path 15 .
- the aerofoil 12 extends radially between and interconnects the outer platform 16 and the inner platform 18 as shown in FIG. 2 .
- the trailing-edge passage 26 extends radially entirely through the outer platform 16 and the inner platform 18 in the illustrative embodiment.
- the aerofoil-shaped passage 24 extends radially entirely through the outer platform 16 and the inner platform 18 .
- the component 10 may be a seal segment or a blade configured to be coupled with a disc of one of the rotating wheel assemblies 22 .
- the aerofoil 12 comprises ceramic materials and the support strut 14 comprises metallic materials as shown in FIG. 3 .
- the aerofoil 12 comprises ceramic matrix composite materials.
- each of the aerofoil 12 and support strut 14 may comprise any suitable materials including ceramics, ceramic matrix composites, metals, alloys, super alloys, etc.
- the gas turbine engine 110 includes a fan 112 , a compressor 114 , a combustor 116 , and a turbine 118 .
- the fan 112 is driven by the turbine 118 and provides thrust for propelling an aircraft as shown in FIG. 1 .
- the compressor 114 compresses and delivers air to the combustor 116 .
- the combustor 116 mixes fuel with the compressed air received from the compressor 114 and ignites the fuel.
- the hot, high pressure products of the combustion reaction in the combustor 116 are directed into the turbine 118 to cause the turbine 118 to rotate about an axis 11 of the gas turbine engine 110 and drive the compressor 114 and the fan 112 .
- the fan 112 may be omitted and/or the turbine 118 drives a propeller, drive shaft, or other suitable alternative.
- the turbine 118 includes a plurality of static turbine vane rings 20 that are fixed relative to the axis 11 and a plurality of rotating wheel assemblies 22 as suggested in FIG. 1 .
- Each turbine vane ring 20 includes a plurality of the vanes 40 .
- the hot gases are conducted through the gas path 15 and interact with the wheel assemblies 22 to rotate the wheel assemblies 22 about the axis 11 .
- the turbine vane rings 20 are positioned to direct the gases toward the wheel assemblies 22 with a desired orientation.
- the components 10 may be seal segments or blades included in the wheel assemblies 22 .
- the component 10 includes the aerofoil 12 as shown in FIG. 3 .
- the aerofoil 12 includes a leading edge 28 , a trailing edge 30 , a pressure side 32 , and a suction side 34 as shown in FIG. 3 .
- the leading edge 28 is located at a foremost position of the aerofoil 12 .
- the trailing edge 30 is spaced apart axially from the leading edge 28 relative to the axis 11 .
- the pressure side 32 extends between and interconnects the leading edge 28 and the trailing edge 30 .
- the suction side extends between and interconnects the leading edge 28 and the trailing edge 30 .
- the aerofoil 12 is formed to define the passage 24 which is aerofoil shaped in the illustrative embodiment as shown in FIG. 3 .
- the passage 24 extends radially at least partway into the aerofoil 12 .
- the passage 24 extends radially through the entire aerofoil 12 , outer platform 16 , and inner platform 18 as shown in FIG. 5 . Cooling fluid may be directed into the passage 24 during use of the gas turbine engine 110 .
- load pads 70 are located in the passage 24 and engage the aerofoil 12 and the support strut 14 to transmit loads from the aerofoil 12 to the support strut 14 .
- the aerofoil-shaped passage 24 has a fore end located adjacent the leading edge 28 and an aft end located toward the trailing edge 30 .
- the aft end of the aerofoil-shaped passage 24 is curved and has a minimum radius 42 .
- the aft end of the aerofoil-shaped passage 24 including the minimum radius, has design restraints reduced along with a reduction of restraints on the size and shape of the trailing edge 30 and the trailing-edge passage 26 .
- the aerofoil 12 if formed to define the trailing-edge passage 26 as shown in FIGS. 3 and 4 .
- the trailing-edge passage 26 extends radially through the entire aerofoil 12 as shown in FIG. 5 .
- the trailing-edge passage 26 extends entirely through the aerofoil 12 , the outer platform 16 , and the inner platform 18 .
- the trailing-edge passage 26 is curvilinear and has a minimum radius 44 .
- the minimum radius 44 of the trailing-edge passage 26 is less than the minimum radius 42 of the aft end of the passage 24 .
- Forming the trailing-edge passage 26 as a separate passage from the passage 24 allows for the different radii to be used with each passage 24 , 26 .
- the trailing-edge passage 26 and the aerofoil-shaped passage 24 are the only voids in the aerofoil 12 .
- the aerofoil 12 includes an outer surface 46 , an inner surface 48 , and an inner surface 50 as shown in FIG. 3 .
- the outer surface 46 is an outermost surface of the aerofoil 12 and defines a portion of the gas path 15 and is exposed to gases in the gas path 15 during use of the gas turbine engine 110 .
- the inner surface 48 defines the aerofoil-shaped passage 24 .
- the inner surface 50 defines the trailing-edge passage 26 .
- the inner surface 50 is continuous in the illustrative embodiment as shown in FIGS. 4 and 5 .
- the inner surface 50 is formed without holes or other passages.
- the inner surface 50 is not formed with axially extending passages that fluidly connect the trailing-edge passage 26 with the gas path 15 or with the passage 24 .
- the inner surface 50 may be machined, formed, and/or textured in some embodiments to obtain a desired heat transfer coefficient to transfer heat from the aerofoil 12 to the cooling fluid directed through the trailing-edge passage 26 .
- the aerofoil 12 includes a core body 54 , a trailing edge filler 56 , and an outer layer 58 as shown in FIG. 3 .
- the core body 54 , the trailing edge filler 56 , and the outer layer 58 are infiltrated together and integrally formed such that the ceramic matrix composite aerofoil 12 is a single, one-piece member.
- the aerofoil 12 is formed from one preform as shown in FIGS. 6 and 7 .
- the core body 54 is made using a core body preform
- the trailing edge filler 56 is made using a trailing edge filler preform
- the outer layer 58 is made using an outer layer preform.
- the preforms are infiltrated together to provide the core body 54 , the trailing edge filler 56 , and the outer layer 58 .
- the core body 54 defines the leading edge 28 , a fore portion of the pressure side 32 , and a fore portion of the suction side 34 of the aerofoil 12 .
- the trailing edge filler 56 defines the trailing edge 30 , an aft portion of the pressure side 32 , and an aft portion of the suction side 34 of the aerofoil 12 .
- the outer layer 58 extends around the core body 54 and the trailing edge filler 56 to provide an outermost surface 46 of the aerofoil 12 .
- the trailing edge filler 56 is made of ceramic matrix composite materials in the illustrative embodiment and is formed to define the entire trailing-edge passage 26 as shown in FIG. 4 . Forming the trailing-edge passage 26 in the trailing edge filler 56 and forming the trailing edge filler 56 separate from the core body 54 using the preforms may reduce design constraints on the core body 54 , the trailing-edge passage 26 , and/or the passage 24 .
- the core body 54 and the outer layer 58 are made of ceramic matrix composite materials and the core body 54 defines the entire aerofoil-shaped passage 24 in the illustrative embodiment.
- Each passage 24 , 26 may be formed in the preform and then set during infiltration. Each passage 24 , 26 may also be formed after the preforms are infiltrated as solid members and then machined into the solid members.
- the outer platform 16 , the inner platform 18 , and the aerofoil 12 are integrally formed from ceramic matrix composite materials. As such, the outer platform 16 , the inner platform 18 , and the aerofoil 12 provide a single, integral, one-piece vane 40 as shown in FIG. 2 . In other embodiments, the outer platform 16 , the inner platform 18 , and the aerofoil 12 may be formed as separate components.
- the component 10 is formed to include an outer brim 62 that extends radially outward from the outer platform 16 and an inner brim 64 that extends radially inward from the inner platform 18 as shown in FIG. 5 .
- the component 10 is located in the turbine 118 of the gas turbine engine 110 and defines a portion of an outer chamber 66 located radially outward of the component 10 and a portion of an inner chamber 68 located radially inward of the component 10 .
- Cooling air may be provided to one or both of the outer chamber 66 and the inner chamber 68 . Cooling air is directed radially through the trailing-edge passage 26 from one of the outer chamber 66 and the inner chamber 68 to the other of the outer chamber 66 and the inner chamber 68 .
- the ceramic matrix composite aerofoil 12 is adapted to withstand high temperatures, but may have relatively low strength compared to the metallic support strut 14 .
- the support strut 14 provides structural strength to the aerofoil 12 by receiving the force loads applied to the aerofoil 12 and transferring them to a casing that surrounds the component 10 .
- the support strut 14 may not be capable of withstanding directly the high temperatures experienced by the aerofoil 12 .
- the support strut 14 is engaged, directly or indirectly via load pads 70 , seals, etc., with the aerofoil 12 to receive force loads from the aerofoil 12 and transfer them to a casing 38 of the engine 110 that is arranged around the turbine 118 as suggested in FIG. 5 .
- the support strut 14 is located in the passage 24 and extends radially through the outer platform 16 , the inner platform 18 , and the aerofoil 12 as shown in FIG. 5 .
- a first end of the support strut 14 is coupled to the casing 38 arranged around the component 10 .
- the second end of the support strut 14 is cantilevered from the casing 38 in the illustrative embodiment.
- the support strut 14 is hollow in the illustrative embodiment.
- the support strut 14 includes holes that extend through the support strut 14 to allow cooling air to pass through the hollow support strut 14 and flow into the passage 24 .
- the support strut 14 is solid.
- a method in accordance with the present disclosure includes a number of steps.
- the method may include providing the aerofoil-shaped core preform and the trailing edge preform.
- the aerofoil-shaped core is formed to define the aerofoil-shaped passage 24 that extends radially there through.
- the trailing edge filler preform is formed to define the trailing-edge passage 26 that extends radially there through.
- the method includes infiltrating the aerofoil-shaped core preform and the trailing edge filler preform together to provide a single, unitary ceramic matrix composite aerofoil component 12 having the aerofoil-shaped passage 24 and the trailing-edge passage 26 .
- the method may include directing a first portion of cooling fluid radially through the aerofoil-shaped passage 24 and directing a second portion of the cooling fluid radially through the trailing-edge passage 26 .
- the aerofoil-shaped passage 24 has a minimum radius and the trailing-edge passage 26 has a minimum radius.
- the minimum radius of the trailing-edge passage 26 is smaller than the minimum radius of the aerofoil-shaped passage 24 in some embodiments.
- the method may include machining the inner surface 50 of the ceramic matrix composite aerofoil 12 that defines the trailing-edge passage 26 to cause the inner surface 50 to have a surface roughness that provides a desired heat transfer between the second portion of the cooling fluid and the inner surface 50 .
- the method may include directing the second portion of the cooling fluid from the outer chamber 66 located radially outward of the ceramic matrix composite aerofoil 12 , radially inward through the trailing-edge passage 26 , and into the inner chamber 68 located radially inward of the ceramic matrix composite aerofoil 12 .
- the method may include directing the second portion of the cooling fluid from the inner chamber 68 located radially inward of the ceramic matrix composite aerofoil 12 , radially outward through the trailing-edge passage 26 , and into the outer chamber 66 located radially outward of the ceramic matrix composite aerofoil 12 .
- the hottest part of a ceramic matrix composite vane 40 or other aerofoil may be at the trailing edge 30 .
- the trailing edge of an aerofoil may be cooled by axial cooling holes formed in the aerofoil where the cooling air is expelled to the gas path. This may provide design constraints associated with: film ejection (pressure, Mach No. location and rate), trailing edge thickness, trailing edge length, and position of the internal cavity (aerofoil-shaped passage 24 ).
- the present disclosure provides aerofoils 12 with cooling holes 26 in the radial direction.
- the cooling air is not being expelled directly to the gas path 15 .
- the present disclosure may allow for a relaxation or removal in the design constraints such as film ejection (pressure, Mach no. location, rate), trailing edge thickness, trailing edge length and position of the internal cavity 24 .
- Some high pressure stage two turbine nozzle guide vanes 40 are made from metallic super-alloys which may be easier to manufacture into a desired shape than ceramic matrix composite vanes. As such, the metallic vanes may not have the same issues with design constraints as compared to ceramic matrix composite vanes.
- Typical ceramic matrix composite cooling holes connect the internal passage with the aerofoil pressure side as per nozzle guide vane convention.
- Ceramic matrix composite components may feature excellent properties when subjected to high temperatures compared to the conventional materials of choice for gas turbine engines such as Nickel based super alloys.
- the benefit of ceramic matrix composite components may allow for a reduction in cooling air flow to be used to cool the components, resulting in an increase in thermal efficiency and, thus, improving specific fuel capacity of the gas turbine engine.
- the present disclosure details how the cooling scheme at the trailing edge 30 could be made to flow in a radial direction to allow a relaxation of design constraints. This could also allow the cooling air to further be used for sealing purposes.
- the present disclosure provides methods and features for cooling the trailing edge 30 of the ceramic matrix composite (CMC) aerofoils 12 such as vanes 40 and blades. This involves conducting the direction of the cooling air in a radial direction instead of axially which may remove or relax design constraints associated with the trailing edge 30 of said aerofoils.
- CMC ceramic matrix composite
- the region of the gas turbine engine 110 that may be hot enough to warrant the use of ceramic matrix composite materials and not too hot to overheat the materials is the high pressure stage two (HP2) of the turbines 118 .
- the materials may be used in blades, seal segments, and nozzle guide vanes (NGVs).
- FIGS. 4, 6, and 7 Examples of shapes of the radial hole(s) to provide cooling to the trailing edge 30 are shown in FIGS. 4, 6, and 7 .
- the surface roughness of the surface 50 defining the cooling hole(s) are altered in some embodiments to obtain a suitable heat transfer coefficient. Changes to the surface roughness in the trailing edge cavity 26 may allow for an altered heat transfer coefficient.
- the radial cooling flow could go from either outer to inner radius or inner to outer radius, depending on engine architecture.
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- General Engineering & Computer Science (AREA)
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Abstract
A component adapted for use in a gas turbine engine includes an aerofoil configured to interact with gases flowing through the gas turbine engine along a gas path. The aerofoil is formed to include a first passage that extends radially at least partway into the aerofoil and a second passage that extends radially into the aerofoil at a trailing edge of the aerofoil.
Description
- The present disclosure relates generally to components for gas turbine engines, and more specifically to aerofoils that comprise ceramic-containing materials.
- Gas turbine engines are used to power aircraft, watercraft, power generators, and the like. Gas turbine engines typically include a compressor, a combustor, and a turbine. The compressor compresses air drawn into the engine and delivers high pressure air to the combustor. In the combustor, fuel is mixed with the high pressure air and is ignited. Products of the combustion reaction in the combustor are directed into the turbine where work is extracted to drive the compressor and, sometimes, an output shaft. Left-over products of the combustion are exhausted out of the turbine and may provide thrust in some applications.
- Products of the combustion reaction directed into the turbine flow over aerofoils included in stationary vanes and rotating blades of the turbine. The interaction of combustion products with the aerofoils heats aerofoils to temperatures that require the aerofoils to be made from high-temperature resistant materials and/or to be actively cooled by supplying relatively cool air to the vanes and blades. To this end, some aerofoils for vanes and blades incorporate composite materials adapted to withstand very high temperatures. Design and manufacture of vanes and blades from composite materials presents challenges because of the geometry and strength required for the parts.
- The present disclosure may comprise one or more of the following features and combinations thereof.
- A component for a gas turbine engine includes a ceramic matrix composite aerofoil. The ceramic matrix composite aerofoil is adapted to conduct gases flowing through a gas path of a gas turbine engine around the component during use of the component. The ceramic matrix composite aerofoil includes a leading edge, a trailing edge spaced apart axially from the leading edge relative to an axis, a pressure side that extends between and interconnects the leading edge and the trailing edge, and a suction side that extends between and interconnects the leading edge and the trailing edge. The ceramic matrix composite aerofoil may be formed to define an aerofoil-shaped passage that extends radially at least partway into the ceramic matrix composite aerofoil.
- The trailing edge, an aft portion of the pressure side, and an aft portion of the suction side of the ceramic matrix composite aerofoil may define a trailing-edge passage. The trailing-edge passage may extend radially through the ceramic matrix composite aerofoil to conduct cooling fluid radially through the trailing edge of the ceramic matrix composite aerofoil.
- In some embodiments, a minimum radius of the trailing-edge passage may be smaller than a minimum radius of an aft end of the aerofoil-shaped passage. In some embodiments, the trailing edge, the aft portion of the pressure side, and the aft portion of the suction side of the ceramic matrix composite aerofoil may be formed without passages that provide fluid communication between the gas path and the trailing-edge passage.
- In some embodiments, the trailing-edge passage may extend radially through the entire component. The ceramic matrix composite aerofoil may include an inner surface that forms the trailing-edge passage and the inner surface may be continuous and without holes. In some embodiments, the trailing-edge passage and the aerofoil-shaped passage are the only voids in the ceramic matrix composite aerofoil.
- In some embodiments, the component may include a ceramic matrix composite outer platform that defines an outer boundary of the gas path and a ceramic matrix composite inner platform spaced apart radially from the ceramic matrix composite outer platform relative to the axis to define an inner boundary of the gas path. The ceramic matrix composite aerofoil may extend radially between the ceramic matrix composite outer platform and the ceramic matrix composite inner platform. The trailing-edge passage may extend radially entirely through the ceramic matrix composite outer platform and the ceramic matrix composite inner platform.
- In some embodiments, the aerofoil-shaped passage extends radially entirely through the ceramic matrix composite outer platform and the inner platform. In some embodiments, the component may include a metallic support strut located in the aero-foil shaped passage of the ceramic matrix composite aerofoil and configured to receive force loads acting on the ceramic matrix composite aerofoil during use of the component.
- In some embodiments, the ceramic matrix composite aerofoil includes a core body, a trailing edge filler, and an outer layer. The core body may define the leading edge, a fore portion of the pressure side, and a fore portion of the suction side of the ceramic matrix composite aerofoil. The trailing edge filler may define the trailing edge, the aft portion of the pressure side, and the aft portion of the suction side of the ceramic matrix composite aerofoil. The outer layer may extend around the core body and the trailing edge filler to provide an outermost surface of the ceramic matrix composite aerofoil.
- In some embodiments, the trailing edge filler may be made of ceramic matrix composite materials. The trailing edge fill may be formed to define the entire trailing-edge passage.
- In some embodiments, the core body and the outer layer may be made of ceramic matrix composite materials. The core body may define the entire aerofoil-shaped passage.
- In some embodiments, the core body, the trailing edge filler, and the outer layer may be integrally formed such that the ceramic matrix composite aerofoil is a single, one-piece member.
- In some embodiments, the ceramic matrix composite aerofoil may be formed to define a plurality of trailing-edge passages that includes the trailing-edge passage. The plurality of trailing-edge passages may include at least a circular shaped trailing-edge passage and an ellipse shaped trailing-edge passage.
- In some embodiments, the trailing-edge passage may be ellipse shaped. In some embodiments, the trailing-edge passage may be generally triangular.
- A method according to an aspect of the present disclosure may include a number of steps. The steps may include providing an aerofoil-shaped core preform that is formed to define an aerofoil-shaped passage that extends radially there through, providing a trailing edge filler preform that is formed to define a trailing-edge passage that extends radially there through, infiltrating the aerofoil-shaped core preform and the trailing edge filler preform together to provide a single, unitary ceramic matrix composite aerofoil component having the aerofoil-shaped passage and the trailing-edge passage, and directing a first portion of a cooling fluid radially through the aerofoil-shaped passage and directing a second portion of the cooling fluid radially through the trailing-edge passage.
- In some embodiments, the aerofoil-shaped passage may have a minimum radius. The trailing-edge passage may have a minimum radius. The minimum radius of the trailing-edge passage may be smaller than the minimum radius of the aerofoil-shaped passage.
- In some embodiments, the method may include directing the second portion of the cooling fluid from an outer chamber located radially outward of the ceramic matrix composite aerofoil component, radially inward through the trailing-edge passage, and into an inner chamber located radially inward of the ceramic matrix composite aerofoil.
- In some embodiments, the method may include directing the second portion of the cooling fluid from an inner chamber located radially inward of the ceramic matrix composite aerofoil component, radially outward through the trailing-edge passage, and into an outer chamber located radially outward of the ceramic matrix composite aerofoil.
- In some embodiments, the method may include machining an inner surface of the ceramic matrix composite aerofoil component that defines the trailing-edge passage to cause the inner surface to have a surface roughness that provides a desired heat transfer between the second portion of the cooling fluid and the inner surface.
- These and other features of the present disclosure will become more apparent from the following description of the illustrative embodiments.
-
FIG. 1 is a cutaway view of a gas turbine engine that includes a fan, a compressor, a combustor, and a turbine, the turbine including static turbine vane rings configured to direct air into adjacent rotating wheel assemblies; -
FIG. 2 is a perspective view of a component of the turbine shown inFIG. 1 suggesting that the component includes a ceramic matrix composite aerofoil formed to include an aerofoil-shaped passage and a trailing-edge passage that extends radially through the trailing edge of the ceramic matrix composite aerofoil; -
FIG. 3 is a section view taken along line 3-3 of the component ofFIG. 2 showing that the component further includes a support strut located in the aerofoil-shaped passage and the ceramic matrix composite aerofoil is made of an aerofoil core, a trailing edge filler formed to define the trailing-edge passage, and an outer layer arranged around the aerofoil core and the trailing edge filler; -
FIG. 4 is an enlarged view of the trailing edge of the component ofFIG. 3 showing an aft end of the aerofoil-shaped passage and the trailing-edge passage formed in the component; -
FIG. 5 is a section view of the turbine of the gas turbine engine ofFIG. 1 showing the component in the turbine; -
FIG. 6 is a section view of a trailing edge of another embodiment of a ceramic matrix composite aerofoil in accordance with the present disclosure showing that the trailing-edge passage is generally triangular; and -
FIG. 7 is a section view of a trailing edge of another embodiment of a ceramic matrix composite aerofoil in accordance with the present disclosure showing that the trailing edge is formed to define a plurality of trailing-edge passages therein. - For the purposes of promoting an understanding of the principles of the disclosure, reference will now be made to a number of illustrative embodiments illustrated in the drawings and specific language will be used to describe the same.
- A
component 10 for use in agas turbine engine 110 is shown inFIG. 2 . Theillustrative component 10 includes a ceramicmatrix composite aerofoil 12 as shown inFIGS. 2-5 . The ceramicmatrix composite aerofoil 12 interacts with hot gases conducted through agas path 15 of thegas turbine engine 110 and conducts the hot gases around thecomponent 10 toward arotating wheel assembly 22 located downstream of thecomponent 10. In other embodiments, thecomponent 10 is a blade included in therotating wheel assembly 22, a seal segment, or other suitable component. - The
aerofoil 12 is formed to include a passage 24 (sometimes called an internal cavity) and a trailing-edge passage 26 as shown inFIGS. 2 and 3 . Thepassage 24 extends radially at least partway into theaerofoil 12 and, illustratively, extends radially through theaerofoil 12. The trailing-edge passage 26 is located at a trailingedge 30 of theaerofoil 12 and extends radially through theaerofoil 12 to provide cooling to the trailingedge 30. In contrast, some gas turbine engines have aerofoils with trailing edges that are cooled by axially extending cooling holes formed in the aerofoils and the cooling air is expelled to the gas path. - The trailing-
edge passage 26 of the present disclosure extends radially through theaerofoil 12 to reduce or remove the design constraints of theaerofoil 12 such as, for example, film air ejection (pressure, Mach no. location, rate) at the trailingedge 30, trailing edge thickness, trailing edge length, and size and position of thepassage 24. Illustratively, the trailing-edge passage 26 is formed in a preform separate from the preform that defines thepassage 24. The trailing-edge passage 26 is not fluidly connected directly with thepassage 24 via other holes or passages. The trailing-edge passage 26 is not fluidly connected with any axially extending holes that are open to thegas path 15. - The
component 10 is avane 40 in the illustrative embodiment and includes ametallic support strut 14, a ceramic matrix compositeouter platform 16, and a ceramic matrix compositeinner platform 18 as shown inFIG. 2 . Themetallic support strut 14 extends through thepassage 24 of the ceramicmatrix composite aerofoil 12 and receives force loads applied to the ceramicmatrix composite aerofoil 12 by the hot gases. Theouter platform 16 defines an outer boundary of thegas path 15 and the inner platform is spaced apart radially from theouter platform 16 and defines an inner boundary of thegas path 15. - The
aerofoil 12 extends radially between and interconnects theouter platform 16 and theinner platform 18 as shown inFIG. 2 . The trailing-edge passage 26 extends radially entirely through theouter platform 16 and theinner platform 18 in the illustrative embodiment. Illustratively, the aerofoil-shapedpassage 24 extends radially entirely through theouter platform 16 and theinner platform 18. In other embodiments, thecomponent 10 may be a seal segment or a blade configured to be coupled with a disc of one of therotating wheel assemblies 22. - In some embodiments, the
aerofoil 12 comprises ceramic materials and thesupport strut 14 comprises metallic materials as shown inFIG. 3 . In the illustrative embodiment, theaerofoil 12 comprises ceramic matrix composite materials. In other embodiments, each of theaerofoil 12 andsupport strut 14 may comprise any suitable materials including ceramics, ceramic matrix composites, metals, alloys, super alloys, etc. - The
gas turbine engine 110 includes afan 112, acompressor 114, acombustor 116, and aturbine 118. Thefan 112 is driven by theturbine 118 and provides thrust for propelling an aircraft as shown inFIG. 1 . Thecompressor 114 compresses and delivers air to thecombustor 116. Thecombustor 116 mixes fuel with the compressed air received from thecompressor 114 and ignites the fuel. The hot, high pressure products of the combustion reaction in thecombustor 116 are directed into theturbine 118 to cause theturbine 118 to rotate about anaxis 11 of thegas turbine engine 110 and drive thecompressor 114 and thefan 112. In other embodiments, thefan 112 may be omitted and/or theturbine 118 drives a propeller, drive shaft, or other suitable alternative. - The
turbine 118 includes a plurality of static turbine vane rings 20 that are fixed relative to theaxis 11 and a plurality ofrotating wheel assemblies 22 as suggested inFIG. 1 . Eachturbine vane ring 20 includes a plurality of thevanes 40. The hot gases are conducted through thegas path 15 and interact with thewheel assemblies 22 to rotate thewheel assemblies 22 about theaxis 11. The turbine vane rings 20 are positioned to direct the gases toward thewheel assemblies 22 with a desired orientation. In other embodiments, thecomponents 10 may be seal segments or blades included in thewheel assemblies 22. - The
component 10 includes theaerofoil 12 as shown inFIG. 3 . Theaerofoil 12 includes aleading edge 28, a trailingedge 30, apressure side 32, and asuction side 34 as shown inFIG. 3 . The leadingedge 28 is located at a foremost position of theaerofoil 12. The trailingedge 30 is spaced apart axially from the leadingedge 28 relative to theaxis 11. Thepressure side 32 extends between and interconnects the leadingedge 28 and the trailingedge 30. The suction side extends between and interconnects the leadingedge 28 and the trailingedge 30. - The
aerofoil 12 is formed to define thepassage 24 which is aerofoil shaped in the illustrative embodiment as shown inFIG. 3 . Thepassage 24 extends radially at least partway into theaerofoil 12. Illustratively, thepassage 24 extends radially through theentire aerofoil 12,outer platform 16, andinner platform 18 as shown inFIG. 5 . Cooling fluid may be directed into thepassage 24 during use of thegas turbine engine 110. Illustratively,load pads 70 are located in thepassage 24 and engage theaerofoil 12 and thesupport strut 14 to transmit loads from theaerofoil 12 to thesupport strut 14. - The aerofoil-shaped
passage 24 has a fore end located adjacent the leadingedge 28 and an aft end located toward the trailingedge 30. The aft end of the aerofoil-shapedpassage 24 is curved and has aminimum radius 42. By forming the radially extending trailing-edge passage 26 at the trailingedge 30, the aft end of the aerofoil-shapedpassage 24, including the minimum radius, has design restraints reduced along with a reduction of restraints on the size and shape of the trailingedge 30 and the trailing-edge passage 26. - The
aerofoil 12 if formed to define the trailing-edge passage 26 as shown inFIGS. 3 and 4 . The trailing-edge passage 26 extends radially through theentire aerofoil 12 as shown inFIG. 5 . Illustratively, the trailing-edge passage 26 extends entirely through theaerofoil 12, theouter platform 16, and theinner platform 18. The trailing-edge passage 26 is curvilinear and has aminimum radius 44. Theminimum radius 44 of the trailing-edge passage 26 is less than theminimum radius 42 of the aft end of thepassage 24. Forming the trailing-edge passage 26 as a separate passage from thepassage 24 allows for the different radii to be used with eachpassage edge passage 26 and the aerofoil-shapedpassage 24 are the only voids in theaerofoil 12. - The trailing-
edge passage 26 may be circular, ellipse shaped, oval, triangular, rectangular, eccentric, uniquely shaped, or any other suitable alternative. In some embodiments, the trailing-edge passage 26 is ellipse shaped as shown inFIG. 4 . In other embodiments, the trailing-edge passage 26 is generally triangular as shown inFIG. 6 . In some embodiments, the trailing-edge passage 26 is circular as shown inFIG. 7 . In some embodiments, theaerofoil 12 is formed to define a plurality of trailing-edge passages 26 as shown inFIG. 7 . - The
aerofoil 12 includes anouter surface 46, aninner surface 48, and aninner surface 50 as shown inFIG. 3 . Theouter surface 46 is an outermost surface of theaerofoil 12 and defines a portion of thegas path 15 and is exposed to gases in thegas path 15 during use of thegas turbine engine 110. Theinner surface 48 defines the aerofoil-shapedpassage 24. Theinner surface 50 defines the trailing-edge passage 26. - The
inner surface 50 is continuous in the illustrative embodiment as shown inFIGS. 4 and 5 . Theinner surface 50 is formed without holes or other passages. For example, unlike some aerofoils, theinner surface 50 is not formed with axially extending passages that fluidly connect the trailing-edge passage 26 with thegas path 15 or with thepassage 24. Theinner surface 50 may be machined, formed, and/or textured in some embodiments to obtain a desired heat transfer coefficient to transfer heat from theaerofoil 12 to the cooling fluid directed through the trailing-edge passage 26. - The
aerofoil 12 includes acore body 54, a trailingedge filler 56, and anouter layer 58 as shown inFIG. 3 . Illustratively, thecore body 54, the trailingedge filler 56, and theouter layer 58 are infiltrated together and integrally formed such that the ceramicmatrix composite aerofoil 12 is a single, one-piece member. In other embodiments, theaerofoil 12 is formed from one preform as shown inFIGS. 6 and 7 . Thecore body 54 is made using a core body preform, the trailingedge filler 56 is made using a trailing edge filler preform, and theouter layer 58 is made using an outer layer preform. The preforms are infiltrated together to provide thecore body 54, the trailingedge filler 56, and theouter layer 58. - The
core body 54 defines the leadingedge 28, a fore portion of thepressure side 32, and a fore portion of thesuction side 34 of theaerofoil 12. The trailingedge filler 56 defines the trailingedge 30, an aft portion of thepressure side 32, and an aft portion of thesuction side 34 of theaerofoil 12. Theouter layer 58 extends around thecore body 54 and the trailingedge filler 56 to provide anoutermost surface 46 of theaerofoil 12. - The trailing
edge filler 56 is made of ceramic matrix composite materials in the illustrative embodiment and is formed to define the entire trailing-edge passage 26 as shown inFIG. 4 . Forming the trailing-edge passage 26 in the trailingedge filler 56 and forming the trailingedge filler 56 separate from thecore body 54 using the preforms may reduce design constraints on thecore body 54, the trailing-edge passage 26, and/or thepassage 24. Thecore body 54 and theouter layer 58 are made of ceramic matrix composite materials and thecore body 54 defines the entire aerofoil-shapedpassage 24 in the illustrative embodiment. - Each
passage passage - In the illustrative embodiment, the
outer platform 16, theinner platform 18, and theaerofoil 12 are integrally formed from ceramic matrix composite materials. As such, theouter platform 16, theinner platform 18, and theaerofoil 12 provide a single, integral, one-piece vane 40 as shown inFIG. 2 . In other embodiments, theouter platform 16, theinner platform 18, and theaerofoil 12 may be formed as separate components. - In the illustrative embodiment, the
component 10 is formed to include anouter brim 62 that extends radially outward from theouter platform 16 and aninner brim 64 that extends radially inward from theinner platform 18 as shown inFIG. 5 . Thecomponent 10 is located in theturbine 118 of thegas turbine engine 110 and defines a portion of anouter chamber 66 located radially outward of thecomponent 10 and a portion of aninner chamber 68 located radially inward of thecomponent 10. Cooling air may be provided to one or both of theouter chamber 66 and theinner chamber 68. Cooling air is directed radially through the trailing-edge passage 26 from one of theouter chamber 66 and theinner chamber 68 to the other of theouter chamber 66 and theinner chamber 68. - The ceramic
matrix composite aerofoil 12 is adapted to withstand high temperatures, but may have relatively low strength compared to themetallic support strut 14. Thesupport strut 14 provides structural strength to theaerofoil 12 by receiving the force loads applied to theaerofoil 12 and transferring them to a casing that surrounds thecomponent 10. Thesupport strut 14 may not be capable of withstanding directly the high temperatures experienced by theaerofoil 12. - The
support strut 14 is engaged, directly or indirectly viaload pads 70, seals, etc., with theaerofoil 12 to receive force loads from theaerofoil 12 and transfer them to acasing 38 of theengine 110 that is arranged around theturbine 118 as suggested inFIG. 5 . Thesupport strut 14 is located in thepassage 24 and extends radially through theouter platform 16, theinner platform 18, and theaerofoil 12 as shown inFIG. 5 . - A first end of the
support strut 14 is coupled to thecasing 38 arranged around thecomponent 10. The second end of thesupport strut 14 is cantilevered from thecasing 38 in the illustrative embodiment. Thesupport strut 14 is hollow in the illustrative embodiment. In some embodiments, thesupport strut 14 includes holes that extend through thesupport strut 14 to allow cooling air to pass through thehollow support strut 14 and flow into thepassage 24. In other embodiments, thesupport strut 14 is solid. - A method in accordance with the present disclosure includes a number of steps. The method may include providing the aerofoil-shaped core preform and the trailing edge preform. The aerofoil-shaped core is formed to define the aerofoil-shaped
passage 24 that extends radially there through. The trailing edge filler preform is formed to define the trailing-edge passage 26 that extends radially there through. The method includes infiltrating the aerofoil-shaped core preform and the trailing edge filler preform together to provide a single, unitary ceramic matrixcomposite aerofoil component 12 having the aerofoil-shapedpassage 24 and the trailing-edge passage 26. The method may include directing a first portion of cooling fluid radially through the aerofoil-shapedpassage 24 and directing a second portion of the cooling fluid radially through the trailing-edge passage 26. - The aerofoil-shaped
passage 24 has a minimum radius and the trailing-edge passage 26 has a minimum radius. The minimum radius of the trailing-edge passage 26 is smaller than the minimum radius of the aerofoil-shapedpassage 24 in some embodiments. The method may include machining theinner surface 50 of the ceramicmatrix composite aerofoil 12 that defines the trailing-edge passage 26 to cause theinner surface 50 to have a surface roughness that provides a desired heat transfer between the second portion of the cooling fluid and theinner surface 50. - The method may include directing the second portion of the cooling fluid from the
outer chamber 66 located radially outward of the ceramicmatrix composite aerofoil 12, radially inward through the trailing-edge passage 26, and into theinner chamber 68 located radially inward of the ceramicmatrix composite aerofoil 12. The method may include directing the second portion of the cooling fluid from theinner chamber 68 located radially inward of the ceramicmatrix composite aerofoil 12, radially outward through the trailing-edge passage 26, and into theouter chamber 66 located radially outward of the ceramicmatrix composite aerofoil 12. - The hottest part of a ceramic matrix
composite vane 40 or other aerofoil may be at the trailingedge 30. In some gas turbine engines, the trailing edge of an aerofoil may be cooled by axial cooling holes formed in the aerofoil where the cooling air is expelled to the gas path. This may provide design constraints associated with: film ejection (pressure, Mach No. location and rate), trailing edge thickness, trailing edge length, and position of the internal cavity (aerofoil-shaped passage 24). - The present disclosure provides
aerofoils 12 withcooling holes 26 in the radial direction. In illustrative embodiments, the cooling air is not being expelled directly to thegas path 15. As such, the present disclosure may allow for a relaxation or removal in the design constraints such as film ejection (pressure, Mach no. location, rate), trailing edge thickness, trailing edge length and position of theinternal cavity 24. - Some high pressure stage two turbine
nozzle guide vanes 40 are made from metallic super-alloys which may be easier to manufacture into a desired shape than ceramic matrix composite vanes. As such, the metallic vanes may not have the same issues with design constraints as compared to ceramic matrix composite vanes. Typical ceramic matrix composite cooling holes connect the internal passage with the aerofoil pressure side as per nozzle guide vane convention. - Ceramic matrix composite components may feature excellent properties when subjected to high temperatures compared to the conventional materials of choice for gas turbine engines such as Nickel based super alloys. The benefit of ceramic matrix composite components may allow for a reduction in cooling air flow to be used to cool the components, resulting in an increase in thermal efficiency and, thus, improving specific fuel capacity of the gas turbine engine.
- The present disclosure details how the cooling scheme at the trailing
edge 30 could be made to flow in a radial direction to allow a relaxation of design constraints. This could also allow the cooling air to further be used for sealing purposes. The present disclosure provides methods and features for cooling the trailingedge 30 of the ceramic matrix composite (CMC) aerofoils 12 such asvanes 40 and blades. This involves conducting the direction of the cooling air in a radial direction instead of axially which may remove or relax design constraints associated with the trailingedge 30 of said aerofoils. - The region of the
gas turbine engine 110 that may be hot enough to warrant the use of ceramic matrix composite materials and not too hot to overheat the materials is the high pressure stage two (HP2) of theturbines 118. The materials may be used in blades, seal segments, and nozzle guide vanes (NGVs). - Examples of shapes of the radial hole(s) to provide cooling to the trailing
edge 30 are shown inFIGS. 4, 6, and 7 . The surface roughness of thesurface 50 defining the cooling hole(s) are altered in some embodiments to obtain a suitable heat transfer coefficient. Changes to the surface roughness in the trailingedge cavity 26 may allow for an altered heat transfer coefficient. The radial cooling flow could go from either outer to inner radius or inner to outer radius, depending on engine architecture. - While the disclosure has been illustrated and described in detail in the foregoing drawings and description, the same is to be considered as exemplary and not restrictive in character, it being understood that only illustrative embodiments thereof have been shown and described and that all changes and modifications that come within the spirit of the disclosure are desired to be protected.
Claims (20)
1. A component for a gas turbine engine, the component comprising:
a ceramic matrix composite aerofoil adapted to conduct gases flowing through a gas path of the gas turbine engine around the component during use of the component, the ceramic matrix composite aerofoil having a leading edge, a trailing edge spaced apart axially from the leading edge relative to an axis, a pressure side that extends between and interconnects the leading edge and the trailing edge, and a suction side that extends between and interconnects the leading edge and the trailing edge,
the ceramic matrix composite aerofoil formed to define an aerofoil-shaped passage that extends radially at least partway into the ceramic matrix composite aerofoil, and
the trailing edge, an aft portion of the pressure side, and an aft portion of the suction side of the ceramic matrix composite aerofoil define a trailing-edge passage that extends radially through the ceramic matrix composite aerofoil to conduct cooling fluid radially through the trailing edge of the ceramic matrix composite aerofoil.
2. The component of claim 1 , wherein a minimum radius of the trailing-edge passage is smaller than a minimum radius of an aft end of the aerofoil-shaped passage.
3. The component of claim 1 , wherein the trailing edge, the aft portion of the pressure side, and the aft portion of the suction side of the ceramic matrix composite aerofoil are formed without passages that provide fluid communication between the gas path and the trailing-edge passage.
4. The component of claim 1 , wherein the trailing-edge passage extends radially through the entire component, the ceramic matrix composite aerofoil includes an inner surface that forms the trailing-edge passage, and the inner surface is continuous and without holes.
5. The component of claim 1 , wherein the trailing-edge passage and the aerofoil-shaped passage are the only voids in the ceramic matrix composite aerofoil.
6. The component of claim 1 , further comprising a ceramic matrix composite outer platform that defines an outer boundary of the gas path and a ceramic matrix composite inner platform spaced apart radially from the ceramic matrix composite outer platform relative to the axis to define an inner boundary of the gas path, the ceramic matrix composite aerofoil extends radially between the ceramic matrix composite outer platform and the ceramic matrix composite inner platform, and the trailing-edge passage extends radially entirely through the ceramic matrix composite outer platform and the ceramic matrix composite inner platform.
7. The component of claim 6 , wherein the aerofoil-shaped passage extends radially entirely through the ceramic matrix composite outer platform and the inner platform.
8. The component of claim 6 , further comprising a metallic support strut located in the aero-foil shaped passage of the ceramic matrix composite aerofoil and configured to receive force loads acting on the ceramic matrix composite aerofoil during use of the component.
9. The component of claim 1 , wherein the ceramic matrix composite aerofoil includes a core body, a trailing edge filler, and an outer layer, the core body defines the leading edge, a fore portion of the pressure side, and a fore portion of the suction side of the ceramic matrix composite aerofoil, the trailing edge filler defines the trailing edge, the aft portion of the pressure side, and the aft portion of the suction side of the ceramic matrix composite aerofoil, and the outer layer extends around the core body and the trailing edge filler to provide an outermost surface of the ceramic matrix composite aerofoil.
10. The component of claim 9 , wherein the trailing edge filler is made of ceramic matrix composite materials and is formed to define the entire trailing-edge passage.
11. The component of claim 10 , wherein the core body and the outer layer are made of ceramic matrix composite materials and the core body defines the entire aerofoil-shaped passage.
12. The component of claim 9 , wherein the core body, the trailing edge filler, and the outer layer are integrally formed such that the ceramic matrix composite aerofoil is a single, one-piece member.
13. The component of claim 12 , wherein the ceramic matrix composite aerofoil is formed to define a plurality of trailing-edge passages that includes the trailing-edge passage, the plurality of trailing-edge passages includes at least a circular shaped trailing-edge passage and an ellipse shaped trailing-edge passage.
14. The component of claim 12 , wherein the trailing-edge passage is ellipse shaped.
15. The component of claim 12 , wherein the trailing-edge passage is generally triangular.
16. A method comprising
providing an aerofoil-shaped core preform that is formed to define an aerofoil-shaped passage that extends radially there through,
providing a trailing edge filler preform that is formed to define a trailing-edge passage that extends radially there through,
infiltrating the aerofoil-shaped core preform and the trailing edge filler preform together to provide a single, unitary ceramic matrix composite aerofoil component having the aerofoil-shaped passage and the trailing-edge passage, and
directing a first portion of a cooling fluid radially through the aerofoil-shaped passage and directing a second portion of the cooling fluid radially through the trailing-edge passage.
17. The method of claim 16 , wherein the aerofoil-shaped passage has a minimum radius, the trailing-edge passage has a minimum radius, and the minimum radius of the trailing-edge passage is smaller than the minimum radius of the aerofoil-shaped passage.
18. The method of claim 16 , further comprising directing the second portion of the cooling fluid from an outer chamber located radially outward of the ceramic matrix composite aerofoil component, radially inward through the trailing-edge passage, and into an inner chamber located radially inward of the ceramic matrix composite aerofoil.
19. The method of claim 16 , further comprising directing the second portion of the cooling fluid from an inner chamber located radially inward of the ceramic matrix composite aerofoil component, radially outward through the trailing-edge passage, and into an outer chamber located radially outward of the ceramic matrix composite aerofoil.
20. The method of claim 16 , further comprising machining an inner surface of the ceramic matrix composite aerofoil component that defines the trailing-edge passage to cause the inner surface to have a surface roughness that provides a desired heat transfer between the second portion of the cooling fluid and the inner surface.
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US20220136393A1 (en) * | 2020-11-02 | 2022-05-05 | Raytheon Technologies Corporation | Cmc vane arc segment with cantilevered spar |
Family Cites Families (18)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US5813827A (en) | 1997-04-15 | 1998-09-29 | Westinghouse Electric Corporation | Apparatus for cooling a gas turbine airfoil |
US6132169A (en) | 1998-12-18 | 2000-10-17 | General Electric Company | Turbine airfoil and methods for airfoil cooling |
US6514046B1 (en) * | 2000-09-29 | 2003-02-04 | Siemens Westinghouse Power Corporation | Ceramic composite vane with metallic substructure |
US7198458B2 (en) | 2004-12-02 | 2007-04-03 | Siemens Power Generation, Inc. | Fail safe cooling system for turbine vanes |
US7527474B1 (en) * | 2006-08-11 | 2009-05-05 | Florida Turbine Technologies, Inc. | Turbine airfoil with mini-serpentine cooling passages |
US7722326B2 (en) | 2007-03-13 | 2010-05-25 | Siemens Energy, Inc. | Intensively cooled trailing edge of thin airfoils for turbine engines |
US8210803B2 (en) * | 2007-06-28 | 2012-07-03 | United Technologies Corporation | Ceramic matrix composite turbine engine vane |
US8206098B2 (en) * | 2007-06-28 | 2012-06-26 | United Technologies Corporation | Ceramic matrix composite turbine engine vane |
US8292580B2 (en) * | 2008-09-18 | 2012-10-23 | Siemens Energy, Inc. | CMC vane assembly apparatus and method |
US8251660B1 (en) | 2009-10-26 | 2012-08-28 | Florida Turbine Technologies, Inc. | Turbine airfoil with near wall vortex cooling |
US8523524B2 (en) | 2010-03-25 | 2013-09-03 | General Electric Company | Airfoil cooling hole flag region |
US8807944B2 (en) | 2011-01-03 | 2014-08-19 | General Electric Company | Turbomachine airfoil component and cooling method therefor |
US8979477B2 (en) | 2011-03-09 | 2015-03-17 | General Electric Company | System for cooling and purging exhaust section of gas turbine engine |
EP2938859B1 (en) | 2012-12-29 | 2019-05-22 | United Technologies Corporation | Cooling architecture for turbine exhaust case |
US9863260B2 (en) * | 2015-03-30 | 2018-01-09 | General Electric Company | Hybrid nozzle segment assemblies for a gas turbine engine |
CA2954785A1 (en) | 2016-01-25 | 2017-07-25 | Rolls-Royce Corporation | Forward flowing serpentine vane |
US10794289B2 (en) * | 2016-08-09 | 2020-10-06 | General Electric Company | Modulated turbine component cooling |
US20180230815A1 (en) | 2017-02-15 | 2018-08-16 | Florida Turbine Technologies, Inc. | Turbine airfoil with thin trailing edge cooling circuit |
-
2019
- 2019-06-21 US US16/448,137 patent/US10883371B1/en active Active
Cited By (2)
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US20220136393A1 (en) * | 2020-11-02 | 2022-05-05 | Raytheon Technologies Corporation | Cmc vane arc segment with cantilevered spar |
US11448075B2 (en) * | 2020-11-02 | 2022-09-20 | Raytheon Technologies Corporation | CMC vane arc segment with cantilevered spar |
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