US20140003937A1 - Component and a method of cooling a component - Google Patents
Component and a method of cooling a component Download PDFInfo
- Publication number
- US20140003937A1 US20140003937A1 US13/826,976 US201313826976A US2014003937A1 US 20140003937 A1 US20140003937 A1 US 20140003937A1 US 201313826976 A US201313826976 A US 201313826976A US 2014003937 A1 US2014003937 A1 US 2014003937A1
- Authority
- US
- United States
- Prior art keywords
- component
- adjacent
- diffusion
- zone
- trailing edge
- 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.)
- Abandoned
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
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
- F01D5/186—Film cooling
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
-
- 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
- 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
- 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
- F05D2240/304—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor related to the trailing edge of a rotor blade
-
- 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
- F05D2250/00—Geometry
- F05D2250/50—Inlet or outlet
- F05D2250/52—Outlet
-
- 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
Definitions
- the present invention relates generally to turbines. More specifically, to a component and a method of cooling a component in turbine.
- One strategy for alleviating thermal stresses is through cooling the airfoils such that the temperatures experienced by the airfoils are lower than that of the hot-gas path. Effective cooling may, for example, allow the airfoils to withstand higher firing temperatures, withstand greater mechanical stresses at high operating temperatures, and/or extend the part-life of the airfoil, all of which may allow the turbine engine to be more cost-effective and efficient.
- One way to cool airfoils during operation is through the use of internal cooling passageways or circuits. Generally, this involves passing a relatively cool supply of compressed air, which may be supplied by the compressor of the turbine engine, through internal cooling circuits within the airfoils. As the compressed air passes through the airfoil, it convectively cools the airfoil, which may allow the part to withstand firing temperatures that it otherwise could not.
- the supply of compressed air is released through small holes on the surface of the airfoils. Released in this manner, the supply of air forms a thin layer or film of relatively cool air at the surface of the airfoil, which both cools and insulates the part from the higher temperatures that surround it.
- This type of cooling which is commonly referred to as “film cooling,” however, comes at an expense.
- film cooling comes at an expense.
- the release of the compressed air in this manner over the surface of the airfoil lowers the aero-efficiency of the engine. As a result, there is an ongoing need for improved cooling strategies for turbine airfoils.
- a component may include a leading edge, a trailing edge, at least one cavity between the leading edge and the trailing edge and at least one diffusion member adjacent to the cavity.
- the diffusion member may include an inlet adjacent to the cavity, a metering zone adjacent to the inlet, a diffusion zone adjacent to the metering zone, and an outlet adjacent the diffusion zone and adjacent the trailing edge.
- the diffusion member may provide up to about 70% reduction in flow and uniform cooling of the trailing edge of the component.
- a method of cooling a component may include providing the component.
- the component may include a leading edge, a trailing edge, at least one cavity between the leading edge and the trailing edge and at least one diffusion member adjacent to the cavity.
- the diffusion member may include an inlet adjacent to the cavity, a metering zone adjacent to the inlet, a diffusion zone adjacent to the metering zone, and an outlet adjacent the diffusion zone and adjacent the trailing edge.
- the diffusion member may provide up to about 70% reduction in flow and uniform cooling of the trailing edge of the component.
- the method may include circulating cooling air in the at least one cavity through the diffusion member. The heat from the component may be removed through the diffusion zone.
- FIG. 1 is a perspective schematic of a component of the present disclosure.
- FIG. 2 is a blown-up view of FIG. 1 of the diffusion zone of the present disclosure.
- FIG. 3 section view of FIG. 1 taken along line 3 - 3 of the present disclosure.
- FIG. 4 is a blown-up view of the diffusion zone of FIG. 3 of the present disclosure.
- FIG. 5 is an alternative embodiment of the diffusion zone of FIG. 3 of the present disclosure.
- FIG. 6 is an alternative embodiment of the diffusion zone of the present disclosure.
- a component and a method of cooling a component are provided.
- One advantage of an embodiment of the present disclosure includes reducing parasitic flows from a turbine. Another advantage of an embodiment of the present disclosure includes reducing the discharge velocity of the nozzle trailing edge cooling slots. Yet another advantage of the present disclosure is increased engine efficiency.
- a component including a leading edge, a trailing edge, at least one cavity between the leading edge and the trailing edge, and at least one diffusion member adjacent to the cavity.
- Component may generally be a hot gas flow path component and may include turbine components, such as, but not limited to, nozzles, blades and shrouds.
- the diffusion member may provide up to about 70% reduction in flow and uniform cooling of the trailing edge of the component.
- the component may be a ceramic matrix composite.
- the component may be a superalloy metal, such as but not limited to nickel-based superalloy, cobalt-based superalloy, or a combination thereof.
- FIG. 1 is a perspective schematic of a component 100 .
- component 100 may be a nozzle.
- Component 100 may include an airfoil 102 .
- Airfoil 102 may be a member between the inner and outer turbine flow path with purpose to change the flow gas path direction. Airfoil may include a leading edge 110 , a trailing edge 112 , and a body 104 between leading edge 110 and trailing edge 112 .
- component 100 may include at least one cavity 200 , 310 , 320 between leading edge 110 and trailing edge 112 .
- Component 100 may include at least one diffusion member 130 adjacent to the at least one cavity.
- diffusion member may include an inlet adjacent to the cavity, a metering zone adjacent to the inlet, a diffusion zone adjacent to the metering zone, and an outlet adjacent the diffusion zone and adjacent the trailing edge.
- diffusion member 130 may include an inlet 210 adjacent to cavity 200 .
- Diffusion member 130 may include a metering zone 220 adjacent to inlet 210 .
- Diffusion member 130 may include a diffusion zone 230 adjacent to metering zone 220 .
- Diffusion member 130 may include an outlet 240 adjacent diffusion zone 220 and adjacent trailing edge 112 .
- Diffusion member 130 may provide up to about 70% reduction in flow and uniform cooling of trailing edge 112 of component 100 .
- Velocity of flow at outlet 240 may be significantly reduced.
- Diffusion member 130 may expand the cross section area from inlet 210 to outlet 240 .
- flow expansion may be a process in which the mass flux is reduced with increasing through-flow area, or a device that enables said process, such as the diffusion member 120 .
- flow metering may be a process or device that controls the quantity of flow traversing the member containing the metering process or device.
- inlet 210 may be the location that flow enters trailing edge 112 cooling scheme from aft cavity 200 (see FIG. 3 ).
- Inlet 210 may be typically short in length with the ratio of the length to hydraulic diameter being less than about 5.
- Inlet 210 may have unique geometric characteristics intended to reduce its metering characteristics.
- Metering zone 220 may be primarily a controlled geometry feature the size of which has the most significant impact on the flow rate passing through trailing edge 112 cooling scheme.
- Metering zone 220 may have a secondary geometry feature that causes or is intended to cause a reduction in the flow rate. Secondary metering may have a non-negligible impact on flow rate but may not be the flow controlling feature.
- Diffusion zone 230 or expansion region may be a region with through-flow area increase of about 150% to about 500%, between the metering zone 220 and outlet 240 .
- Diffusion zone 230 may include a diffusion angle 232 .
- Diffusion angle 232 may be any angle that provides the desired expansion of flow.
- Outlet 240 may be the location where flow exits internal portion of trailing edge 112 cooling scheme. Outlet 240 may be characterized by film coverage and in the event outlet 240 bisects the trailing edge 112 with no film coverage zone, then outlet 240 may have an open-to-solid ratio in the range of about 25% to about 100%.
- film coverage may be measured in the direction orthogonal to the flow, and is the fraction of the distance that is exposed to outlet 240 .
- FIG. 3 is a sectional view along line 3 - 3 of FIG. 1 and shows forward cavity 320 adjacent leading edge 110 .
- Second cavity 310 may be adjacent forward cavity 320 .
- Aft cavity 200 may be adjacent diffusion member 130 .
- Diffusion member 130 may be adjacent external portion 350 .
- External portion 350 may be the distance between outlet 240 of diffusion member 130 and airfoil closeout, having a ratio of length (L) to hydraulic diameter (D) in the range of about zero to about twelve.
- L length
- D hydraulic diameter
- the film coverage may be in the range of about 33% to 100%.
- the outlet 240 may have “open” coverage in the range of about 33% to 100%.
- FIG. 4 illustrates a schematic blow-up of FIG. 3 highlighting the diffusion member 130 .
- Inlet 210 may be adjacent to aft cavity 200 .
- Metering zone 220 may be adjacent inlet 210 .
- Diffusion zone 230 may be adjacent metering zone 220 .
- Outlet 240 may be adjacent diffusion zone and external portion 350 of airfoil 102 trailing edge 112 .
- outlet 240 may exit at base of trailing edge 112 , having no breakout length.
- a diffusion member may include two or more diffusion zones.
- diffusion member 130 may include inlet 210 adjacent to aft cavity 220 of component 100 .
- Inlet 210 may be adjacent to metering zone 220 .
- There may be two or more diffusion zones 230 adjacent to metering zone 220 .
- Each diffusion zone 230 may provide the desired expansion and decrease in flow.
- Each diffusion zone 230 may include an outlet 240 adjacent to trailing edge 112 .
- Diffusion member 130 may be formed in component 100 using any suitable technologies, such as, but not limited to, lasers, or electrical discharge machining (EDM).
- suitable technologies such as, but not limited to, lasers, or electrical discharge machining (EDM).
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
Description
- This patent application claims the benefit of U.S. Provisional Patent Application Ser. No. 61/666,813 filed on Jun. 30, 2012 and entitled “A COMPONENT AND A METHOD OF COOLING A COMPONENT,” the disclosure of which is incorporated by reference as if fully rewritten herein.
- The present invention relates generally to turbines. More specifically, to a component and a method of cooling a component in turbine.
- The objective of designing and building more efficient turbine engines is a significant one, particularly considering the growing scarcity and increasing cost of fossil fuels. While several strategies for increasing the efficiency of turbine engines are known, it remains a challenging goal because the known alternatives, which, for example, include increasing the size of the engine, increasing the temperatures through the hot-gas path, and increasing the rotational velocities of the rotor blades, generally place additional strain on parts, including additional strain on turbine airfoils, which are already highly stressed. As a result, improved apparatus, methods and/or systems that reduce operational stresses placed on turbine airfoils or allow the turbine airfoils to better withstand these stresses are in great demand.
- One strategy for alleviating thermal stresses is through cooling the airfoils such that the temperatures experienced by the airfoils are lower than that of the hot-gas path. Effective cooling may, for example, allow the airfoils to withstand higher firing temperatures, withstand greater mechanical stresses at high operating temperatures, and/or extend the part-life of the airfoil, all of which may allow the turbine engine to be more cost-effective and efficient. One way to cool airfoils during operation is through the use of internal cooling passageways or circuits. Generally, this involves passing a relatively cool supply of compressed air, which may be supplied by the compressor of the turbine engine, through internal cooling circuits within the airfoils. As the compressed air passes through the airfoil, it convectively cools the airfoil, which may allow the part to withstand firing temperatures that it otherwise could not.
- In some instances, the supply of compressed air is released through small holes on the surface of the airfoils. Released in this manner, the supply of air forms a thin layer or film of relatively cool air at the surface of the airfoil, which both cools and insulates the part from the higher temperatures that surround it. This type of cooling, which is commonly referred to as “film cooling,” however, comes at an expense. The release of the compressed air in this manner over the surface of the airfoil, lowers the aero-efficiency of the engine. As a result, there is an ongoing need for improved cooling strategies for turbine airfoils.
- Therefore, a component and a method of cooling a component in turbine that do not suffer from the above drawbacks is desirable in the art.
- According to an exemplary embodiment of the present disclosure, a component is provided. The component may include a leading edge, a trailing edge, at least one cavity between the leading edge and the trailing edge and at least one diffusion member adjacent to the cavity. The diffusion member may include an inlet adjacent to the cavity, a metering zone adjacent to the inlet, a diffusion zone adjacent to the metering zone, and an outlet adjacent the diffusion zone and adjacent the trailing edge. The diffusion member may provide up to about 70% reduction in flow and uniform cooling of the trailing edge of the component.
- According to another exemplary embodiment of the present disclosure, a method of cooling a component is provided. The method may include providing the component. The component may include a leading edge, a trailing edge, at least one cavity between the leading edge and the trailing edge and at least one diffusion member adjacent to the cavity. The diffusion member may include an inlet adjacent to the cavity, a metering zone adjacent to the inlet, a diffusion zone adjacent to the metering zone, and an outlet adjacent the diffusion zone and adjacent the trailing edge. The diffusion member may provide up to about 70% reduction in flow and uniform cooling of the trailing edge of the component. The method may include circulating cooling air in the at least one cavity through the diffusion member. The heat from the component may be removed through the diffusion zone.
- Other features and advantages of the present invention will be apparent from the following more detailed description of the preferred embodiment, taken in conjunction with the accompanying drawings which illustrate, by way of example, the principles of the invention.
-
FIG. 1 is a perspective schematic of a component of the present disclosure. -
FIG. 2 is a blown-up view ofFIG. 1 of the diffusion zone of the present disclosure. -
FIG. 3 section view ofFIG. 1 taken along line 3-3 of the present disclosure. -
FIG. 4 is a blown-up view of the diffusion zone ofFIG. 3 of the present disclosure. -
FIG. 5 is an alternative embodiment of the diffusion zone ofFIG. 3 of the present disclosure. -
FIG. 6 is an alternative embodiment of the diffusion zone of the present disclosure. - Wherever possible, the same reference numbers will be used throughout the drawings to represent the same parts.
- Provided is a component and a method of cooling a component.
- One advantage of an embodiment of the present disclosure includes reducing parasitic flows from a turbine. Another advantage of an embodiment of the present disclosure includes reducing the discharge velocity of the nozzle trailing edge cooling slots. Yet another advantage of the present disclosure is increased engine efficiency.
- According to one embodiment, a component including a leading edge, a trailing edge, at least one cavity between the leading edge and the trailing edge, and at least one diffusion member adjacent to the cavity is provided. Component may generally be a hot gas flow path component and may include turbine components, such as, but not limited to, nozzles, blades and shrouds. According to one embodiment the diffusion member may provide up to about 70% reduction in flow and uniform cooling of the trailing edge of the component. In one embodiment, the component may be a ceramic matrix composite. In another embodiment, the component may be a superalloy metal, such as but not limited to nickel-based superalloy, cobalt-based superalloy, or a combination thereof.
- For example,
FIG. 1 is a perspective schematic of acomponent 100. For example, as depicted,component 100 may be a nozzle.Component 100 may include anairfoil 102. Airfoil 102 may be a member between the inner and outer turbine flow path with purpose to change the flow gas path direction. Airfoil may include a leadingedge 110, atrailing edge 112, and a body 104 between leadingedge 110 andtrailing edge 112. As shown inFIG. 2 , for example,component 100 may include at least onecavity edge 110 andtrailing edge 112.Component 100 may include at least onediffusion member 130 adjacent to the at least one cavity. - According to one embodiment, diffusion member may include an inlet adjacent to the cavity, a metering zone adjacent to the inlet, a diffusion zone adjacent to the metering zone, and an outlet adjacent the diffusion zone and adjacent the trailing edge. For example, as shown in
FIG. 2 ,diffusion member 130 may include aninlet 210 adjacent tocavity 200.Diffusion member 130 may include ametering zone 220 adjacent toinlet 210.Diffusion member 130 may include adiffusion zone 230 adjacent tometering zone 220.Diffusion member 130 may include anoutlet 240adjacent diffusion zone 220 and adjacenttrailing edge 112.Diffusion member 130 may provide up to about 70% reduction in flow and uniform cooling oftrailing edge 112 ofcomponent 100. Velocity of flow atoutlet 240 may be significantly reduced.Diffusion member 130 may expand the cross section area frominlet 210 tooutlet 240. As used herein “flow expansion” may be a process in which the mass flux is reduced with increasing through-flow area, or a device that enables said process, such as thediffusion member 120. As used herein “flow metering” may be a process or device that controls the quantity of flow traversing the member containing the metering process or device. - As shown in
FIG. 2 ,inlet 210 may be the location that flow enters trailingedge 112 cooling scheme from aft cavity 200 (seeFIG. 3 ).Inlet 210 may be typically short in length with the ratio of the length to hydraulic diameter being less than about 5.Inlet 210 may have unique geometric characteristics intended to reduce its metering characteristics.Metering zone 220 may be primarily a controlled geometry feature the size of which has the most significant impact on the flow rate passing through trailingedge 112 cooling scheme.Metering zone 220 may have a secondary geometry feature that causes or is intended to cause a reduction in the flow rate. Secondary metering may have a non-negligible impact on flow rate but may not be the flow controlling feature. -
Diffusion zone 230 or expansion region may be a region with through-flow area increase of about 150% to about 500%, between themetering zone 220 andoutlet 240.Diffusion zone 230 may include adiffusion angle 232.Diffusion angle 232 may be any angle that provides the desired expansion of flow.Outlet 240 may be the location where flow exits internal portion of trailingedge 112 cooling scheme.Outlet 240 may be characterized by film coverage and in theevent outlet 240 bisects the trailingedge 112 with no film coverage zone, thenoutlet 240 may have an open-to-solid ratio in the range of about 25% to about 100%. As used herein, “film coverage,” may be measured in the direction orthogonal to the flow, and is the fraction of the distance that is exposed tooutlet 240. -
FIG. 3 is a sectional view along line 3-3 ofFIG. 1 and showsforward cavity 320 adjacentleading edge 110.Second cavity 310 may be adjacentforward cavity 320.Aft cavity 200 may beadjacent diffusion member 130.Diffusion member 130 may be adjacentexternal portion 350.External portion 350 may be the distance betweenoutlet 240 ofdiffusion member 130 and airfoil closeout, having a ratio of length (L) to hydraulic diameter (D) in the range of about zero to about twelve. In one embodiment, when the L/D ratio may be finite, the film coverage may be in the range of about 33% to 100%. In another embodiment when the L/D ratio may be zero, theoutlet 240 may have “open” coverage in the range of about 33% to 100%. - According to one embodiment, a diffusion member is provided. For example,
FIG. 4 illustrates a schematic blow-up ofFIG. 3 highlighting thediffusion member 130.Inlet 210 may be adjacent toaft cavity 200.Metering zone 220 may beadjacent inlet 210.Diffusion zone 230 may beadjacent metering zone 220.Outlet 240 may be adjacent diffusion zone andexternal portion 350 ofairfoil 102trailing edge 112. In an alternative embodiment, as shown inFIG. 5 ,outlet 240 may exit at base of trailingedge 112, having no breakout length. - According to one embodiment, a diffusion member may include two or more diffusion zones. For example, as illustrated in
FIG. 6 diffusion member 130 may includeinlet 210 adjacent toaft cavity 220 ofcomponent 100.Inlet 210 may be adjacent tometering zone 220. There may be two ormore diffusion zones 230 adjacent tometering zone 220. Eachdiffusion zone 230 may provide the desired expansion and decrease in flow. Eachdiffusion zone 230 may include anoutlet 240 adjacent to trailingedge 112. -
Diffusion member 130 may be formed incomponent 100 using any suitable technologies, such as, but not limited to, lasers, or electrical discharge machining (EDM). - While the invention has been described with reference to a preferred embodiment, it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the invention. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the invention without departing from the essential scope thereof. Therefore, it is intended that the invention not be limited to the particular embodiment disclosed as the best mode contemplated for carrying out this invention, but that the invention will include all embodiments falling within the scope of the appended claims.
Claims (14)
Priority Applications (7)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US13/826,976 US20140003937A1 (en) | 2012-06-30 | 2013-03-14 | Component and a method of cooling a component |
JP2015520221A JP2015526629A (en) | 2012-06-30 | 2013-06-06 | Parts and parts cooling method |
PCT/US2013/044415 WO2014004014A1 (en) | 2012-06-30 | 2013-06-06 | A component and a method of cooling a component |
CN201380035010.2A CN104379874A (en) | 2012-06-30 | 2013-06-06 | A component and a method of cooling a component |
EP13730991.0A EP2882940A1 (en) | 2012-06-30 | 2013-06-06 | A component and a method of cooling a component |
CA2877330A CA2877330A1 (en) | 2012-06-30 | 2013-06-06 | A component and a method of cooling a component |
BR112015000077A BR112015000077A2 (en) | 2012-06-30 | 2013-06-06 | component and cooling method of a component |
Applications Claiming Priority (2)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US201261666813P | 2012-06-30 | 2012-06-30 | |
US13/826,976 US20140003937A1 (en) | 2012-06-30 | 2013-03-14 | Component and a method of cooling a component |
Publications (1)
Publication Number | Publication Date |
---|---|
US20140003937A1 true US20140003937A1 (en) | 2014-01-02 |
Family
ID=49778353
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
US13/826,976 Abandoned US20140003937A1 (en) | 2012-06-30 | 2013-03-14 | Component and a method of cooling a component |
Country Status (7)
Country | Link |
---|---|
US (1) | US20140003937A1 (en) |
EP (1) | EP2882940A1 (en) |
JP (1) | JP2015526629A (en) |
CN (1) | CN104379874A (en) |
BR (1) | BR112015000077A2 (en) |
CA (1) | CA2877330A1 (en) |
WO (1) | WO2014004014A1 (en) |
Cited By (2)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20170198595A1 (en) * | 2016-01-08 | 2017-07-13 | General Electric Company | Turbine Airfoil Trailing Edge Cooling Passage |
EP4198264A1 (en) * | 2021-12-17 | 2023-06-21 | Raytheon Technologies Corporation | Gas turbine engine component with manifold cavity and metering inlet orifices |
Families Citing this family (4)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US10309227B2 (en) * | 2016-10-26 | 2019-06-04 | General Electric Company | Multi-turn cooling circuits for turbine blades |
US10450950B2 (en) * | 2016-10-26 | 2019-10-22 | General Electric Company | Turbomachine blade with trailing edge cooling circuit |
US10352176B2 (en) * | 2016-10-26 | 2019-07-16 | General Electric Company | Cooling circuits for a multi-wall blade |
US11814965B2 (en) | 2021-11-10 | 2023-11-14 | General Electric Company | Turbomachine blade trailing edge cooling circuit with turn passage having set of obstructions |
Citations (8)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US6176678B1 (en) * | 1998-11-06 | 2001-01-23 | General Electric Company | Apparatus and methods for turbine blade cooling |
US6287075B1 (en) * | 1997-10-22 | 2001-09-11 | General Electric Company | Spanwise fan diffusion hole airfoil |
US20060073016A1 (en) * | 2004-10-04 | 2006-04-06 | Alstom Technology Ltd | Gas turbine airfoil leading edge cooling construction |
US20100068068A1 (en) * | 2008-09-16 | 2010-03-18 | Siemens Energy, Inc. | Turbine Airfoil Cooling System with Diffusion Film Cooling Hole Having Flow Restriction Rib |
US20100074763A1 (en) * | 2008-09-25 | 2010-03-25 | Siemens Energy, Inc. | Trailing Edge Cooling Slot Configuration for a Turbine Airfoil |
US7762775B1 (en) * | 2007-05-31 | 2010-07-27 | Florida Turbine Technologies, Inc. | Turbine airfoil with cooled thin trailing edge |
US20100282721A1 (en) * | 2009-05-05 | 2010-11-11 | General Electric Company | System and method for improved film cooling |
US20130209234A1 (en) * | 2012-02-15 | 2013-08-15 | United Technologies Corporation | Multiple diffusing cooling hole |
Family Cites Families (9)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US4684323A (en) * | 1985-12-23 | 1987-08-04 | United Technologies Corporation | Film cooling passages with curved corners |
JP3302370B2 (en) * | 1995-04-11 | 2002-07-15 | ユナイテッド・テクノロジーズ・コーポレーション | External air seal for turbine blades with thin film cooling slots |
EP1167690A1 (en) * | 2000-06-21 | 2002-01-02 | Siemens Aktiengesellschaft | Cooling of the trailing edge of a gas turbine airfoil |
US6969230B2 (en) * | 2002-12-17 | 2005-11-29 | General Electric Company | Venturi outlet turbine airfoil |
US7374401B2 (en) * | 2005-03-01 | 2008-05-20 | General Electric Company | Bell-shaped fan cooling holes for turbine airfoil |
US7686578B2 (en) * | 2006-08-21 | 2010-03-30 | General Electric Company | Conformal tip baffle airfoil |
US7549844B2 (en) * | 2006-08-24 | 2009-06-23 | Siemens Energy, Inc. | Turbine airfoil cooling system with bifurcated and recessed trailing edge exhaust channels |
US8057181B1 (en) * | 2008-11-07 | 2011-11-15 | Florida Turbine Technologies, Inc. | Multiple expansion film cooling hole for turbine airfoil |
US20110268583A1 (en) * | 2010-04-30 | 2011-11-03 | General Electric Company | Airfoil trailing edge and method of manufacturing the same |
-
2013
- 2013-03-14 US US13/826,976 patent/US20140003937A1/en not_active Abandoned
- 2013-06-06 WO PCT/US2013/044415 patent/WO2014004014A1/en active Application Filing
- 2013-06-06 EP EP13730991.0A patent/EP2882940A1/en not_active Withdrawn
- 2013-06-06 CA CA2877330A patent/CA2877330A1/en not_active Abandoned
- 2013-06-06 BR BR112015000077A patent/BR112015000077A2/en not_active IP Right Cessation
- 2013-06-06 JP JP2015520221A patent/JP2015526629A/en active Pending
- 2013-06-06 CN CN201380035010.2A patent/CN104379874A/en active Pending
Patent Citations (8)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US6287075B1 (en) * | 1997-10-22 | 2001-09-11 | General Electric Company | Spanwise fan diffusion hole airfoil |
US6176678B1 (en) * | 1998-11-06 | 2001-01-23 | General Electric Company | Apparatus and methods for turbine blade cooling |
US20060073016A1 (en) * | 2004-10-04 | 2006-04-06 | Alstom Technology Ltd | Gas turbine airfoil leading edge cooling construction |
US7762775B1 (en) * | 2007-05-31 | 2010-07-27 | Florida Turbine Technologies, Inc. | Turbine airfoil with cooled thin trailing edge |
US20100068068A1 (en) * | 2008-09-16 | 2010-03-18 | Siemens Energy, Inc. | Turbine Airfoil Cooling System with Diffusion Film Cooling Hole Having Flow Restriction Rib |
US20100074763A1 (en) * | 2008-09-25 | 2010-03-25 | Siemens Energy, Inc. | Trailing Edge Cooling Slot Configuration for a Turbine Airfoil |
US20100282721A1 (en) * | 2009-05-05 | 2010-11-11 | General Electric Company | System and method for improved film cooling |
US20130209234A1 (en) * | 2012-02-15 | 2013-08-15 | United Technologies Corporation | Multiple diffusing cooling hole |
Cited By (5)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20170198595A1 (en) * | 2016-01-08 | 2017-07-13 | General Electric Company | Turbine Airfoil Trailing Edge Cooling Passage |
CN106968722A (en) * | 2016-01-08 | 2017-07-21 | 通用电气公司 | Turbine airfoil trailing edge cooling channel |
US10301954B2 (en) * | 2016-01-08 | 2019-05-28 | General Electric Company | Turbine airfoil trailing edge cooling passage |
CN106968722B (en) * | 2016-01-08 | 2021-06-18 | 通用电气公司 | Turbine airfoil trailing edge cooling passage |
EP4198264A1 (en) * | 2021-12-17 | 2023-06-21 | Raytheon Technologies Corporation | Gas turbine engine component with manifold cavity and metering inlet orifices |
Also Published As
Publication number | Publication date |
---|---|
CA2877330A1 (en) | 2014-01-03 |
BR112015000077A2 (en) | 2017-10-10 |
EP2882940A1 (en) | 2015-06-17 |
CN104379874A (en) | 2015-02-25 |
WO2014004014A1 (en) | 2014-01-03 |
JP2015526629A (en) | 2015-09-10 |
Similar Documents
Publication | Publication Date | Title |
---|---|---|
US20140003937A1 (en) | Component and a method of cooling a component | |
US20190085705A1 (en) | Component for a turbine engine with a film-hole | |
US11773729B2 (en) | Component for a gas turbine engine with a film hole | |
EP3495617B1 (en) | Airfoil with internal cooling passages | |
US20200024951A1 (en) | Component for a turbine engine with a cooling hole | |
US11927110B2 (en) | Component for a turbine engine with a cooling hole | |
US20200141247A1 (en) | Component for a turbine engine with a film hole | |
US10760431B2 (en) | Component for a turbine engine with a cooling hole | |
US8105014B2 (en) | Gas turbine engine article having columnar microstructure | |
EP3495615B1 (en) | Airfoil with internal cooling passages | |
EP3495618B1 (en) | Airfoil with internal cooling passages | |
EP3495616B1 (en) | Airfoil with internal cooling passages | |
US7681623B2 (en) | Casting process and cast component | |
EP3495619B1 (en) | Airfoil with internal cooling passages | |
US10626735B2 (en) | Double wall turbine gas turbine engine blade cooling configuration |
Legal Events
Date | Code | Title | Description |
---|---|---|---|
AS | Assignment |
Owner name: GENERAL ELECTRIC COMPANY, NEW YORK Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNORS:GROOMS, JAMES HAMILTON;SINILE, DARRELL GLENN;FREDERICK, ROBERT ALAN;REEL/FRAME:030177/0767 Effective date: 20130313 |
|
AS | Assignment |
Owner name: GENERAL ELECTRIC COMPANY, NEW YORK Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNORS:GROOMS, JAMES HAMILTON, II;SENILE, DARRELL GLENN;FREDERICK, ROBERT ALAN;REEL/FRAME:034531/0135 Effective date: 20140313 |
|
STCB | Information on status: application discontinuation |
Free format text: ABANDONED -- FAILURE TO RESPOND TO AN OFFICE ACTION |