EP2246133B1 - RMC-defined tip blowing slots for turbine blades - Google Patents
RMC-defined tip blowing slots for turbine blades Download PDFInfo
- Publication number
- EP2246133B1 EP2246133B1 EP10007545.6A EP10007545A EP2246133B1 EP 2246133 B1 EP2246133 B1 EP 2246133B1 EP 10007545 A EP10007545 A EP 10007545A EP 2246133 B1 EP2246133 B1 EP 2246133B1
- Authority
- EP
- European Patent Office
- Prior art keywords
- tip
- cooling
- slots
- refractory metal
- metal core
- 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.)
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Links
- 238000007664 blowing Methods 0.000 title description 5
- 238000001816 cooling Methods 0.000 claims description 56
- 239000012809 cooling fluid Substances 0.000 claims 1
- 239000011162 core material Substances 0.000 description 69
- 239000000919 ceramic Substances 0.000 description 31
- 239000003870 refractory metal Substances 0.000 description 30
- 238000005266 casting Methods 0.000 description 21
- 238000000034 method Methods 0.000 description 12
- 239000007787 solid Substances 0.000 description 8
- 230000008901 benefit Effects 0.000 description 6
- 238000005516 engineering process Methods 0.000 description 3
- 239000000853 adhesive Substances 0.000 description 2
- 230000001070 adhesive effect Effects 0.000 description 2
- 238000004891 communication Methods 0.000 description 2
- 230000009429 distress Effects 0.000 description 2
- 230000003628 erosive effect Effects 0.000 description 2
- 239000012530 fluid Substances 0.000 description 2
- 238000005495 investment casting Methods 0.000 description 2
- 230000003647 oxidation Effects 0.000 description 2
- 238000007254 oxidation reaction Methods 0.000 description 2
- 230000000087 stabilizing effect Effects 0.000 description 2
- 229910001182 Mo alloy Inorganic materials 0.000 description 1
- ZOKXTWBITQBERF-UHFFFAOYSA-N Molybdenum Chemical compound [Mo] ZOKXTWBITQBERF-UHFFFAOYSA-N 0.000 description 1
- VYPSYNLAJGMNEJ-UHFFFAOYSA-N Silicium dioxide Chemical group O=[Si]=O VYPSYNLAJGMNEJ-UHFFFAOYSA-N 0.000 description 1
- 239000011248 coating agent Substances 0.000 description 1
- 238000000576 coating method Methods 0.000 description 1
- 230000000694 effects Effects 0.000 description 1
- 238000003754 machining Methods 0.000 description 1
- 238000004519 manufacturing process Methods 0.000 description 1
- 230000000873 masking effect Effects 0.000 description 1
- 229910052751 metal Inorganic materials 0.000 description 1
- 239000002184 metal Substances 0.000 description 1
- 229910052750 molybdenum Inorganic materials 0.000 description 1
- 239000011733 molybdenum Substances 0.000 description 1
- 239000011819 refractory material Substances 0.000 description 1
- 239000000126 substance Substances 0.000 description 1
Images
Classifications
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B22—CASTING; POWDER METALLURGY
- B22C—FOUNDRY MOULDING
- B22C9/00—Moulds or cores; Moulding processes
- B22C9/10—Cores; Manufacture or installation of cores
- B22C9/103—Multipart cores
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B22—CASTING; POWDER METALLURGY
- B22C—FOUNDRY MOULDING
- B22C9/00—Moulds or cores; Moulding processes
- B22C9/02—Sand moulds or like moulds for shaped castings
- B22C9/04—Use of lost patterns
-
- 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
- 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
- F05D2230/00—Manufacture
- F05D2230/20—Manufacture essentially without removing material
- F05D2230/21—Manufacture essentially without removing material by casting
- F05D2230/211—Manufacture essentially without removing material by casting by precision casting, e.g. microfusing or investment casting
-
- 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
- F05D2300/10—Metals, alloys or intermetallic compounds
- F05D2300/13—Refractory metals, i.e. Ti, V, Cr, Zr, Nb, Mo, Hf, Ta, W
-
- 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
- Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
- Y10T—TECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
- Y10T29/00—Metal working
- Y10T29/49—Method of mechanical manufacture
- Y10T29/49316—Impeller making
- Y10T29/49336—Blade making
- Y10T29/49337—Composite blade
Definitions
- the present invention relates to a turbine engine component, such as a turbine blade, having a plurality of as-cast blowing slots in a tip region.
- HPT high pressure turbine
- blade distress via oxidation and erosion. It is particularly challenging to design a cooling configuration for a tip region for a variety of reasons.
- the tip region of a turbine blade is typically the thinnest portion of the airfoil, which makes it more difficult to package the desired cooling features.
- the tip region of a turbine blade is typically difficult to accurately produce with investment casting processes because the internal ceramic core is thin and weak near the tip. Further, it is cantilevered relatively far from the core-locating fixture at the blade root. Considering these points, it is desirable to have methods to create intricate cooling features near the tip capable of being targeted at specific regions of high heat load, while also allowing for greater control during the investment casting process.
- FIG. 1 An existing HPT blade tip cooling design is shown in FIG. 1 .
- a radially oriented cavity supplies cooling air to a leading edge impingement cooling scheme as well as a laterally-oriented cavity, known as a tip flag, that helps cool the tip before exiting the blade at the trailing edge near the tip.
- FIG. 1 also shows a midbody three-pass serpentine cooling arrangement and a trailing edge double-impingement system.
- the tip of the core in FIG. 1 includes an appendage that creates a recess blade tip known as a squealer pocket. That appendage is connected to the leading edge and tip flag core by means of two cylindrical connections (“print-outs") that form open holes in the finished casting (“print-out holes”).
- the core is fixed at the root of the blade during the casting process.
- the squealer pocket core is located laterally during the casting process, allowing the tip print outs to stabilize the tip region of the core. In order to prevent core breakage during the casting process, these tip print-outs should be as large as possible, especially considering that they are constructed from the brittle ceramic core material.
- One of the primary purposes of the squealer pocket is to allow for a shorter distance that the tip print-outs must span.
- EP-1445424-A2 describes an embedded microcircuit cooling design for a turbine blade tip.
- a new tip cooling design that utilizes refractory metal core (RMC) technology in order to create a tip cooling scheme for a turbine engine component that is capable of more efficient use of cooling air and a more reliable casting process.
- RMC refractory metal core
- the process comprises the steps of placing a ceramic core having a configuration of a passageway to be formed in the airfoil portion within a mold; attaching a refractory metal core element to the ceramic core to stabilize a tip region of the ceramic core during casting; and casting the airfoil portion.
- the process may further comprise locating the ceramic core relative to the mold with the refractory metal core element.
- the locating step may comprise providing a refractory metal core element having at least one leg.
- the refractory metal core element providing step may comprise providing a refractory metal core element having a plurality of legs.
- the process may further comprise removing the ceramic core so as to form the passageway and subsequently removing said refractory metal core element and thereby leaving at least one cooling slot in a tip region of the airfoil portion.
- the removing step may comprise leaving a plurality of cooling slots in said tip region.
- the process may further comprise machining a plurality of film cooling holes in the airfoil portion in the vicinity of the passageway formed by the ceramic core.
- a ceramic core for forming a passageway in a cast airfoil portion and means for stabilizing a tip region of the ceramic core.
- the stabilizing means comprises a refractory metal core element.
- the refractory metal core element may comprise a solid portion and a plurality of legs depending from the solid portion. Each leg may have an angled portion and a base portion and the base portion of the legs may be joined together by a lower portion.
- a refractory metal core element comprising a solid portion and a plurality of spaced apart legs depending from the solid portion.
- Each of the legs has a first portion adjacent the solid portion, a base portion, and an angled portion intermediate the first portion and the base portion so that the base portion is laterally offset from the solid portion.
- the base portions of the legs are preferably joined together by a lower portion.
- a new tip cooling design for a turbine blade is proposed here that utilizes refractory metal core technology in order to help create a tip cooling scheme that is capable of more efficient use of cooling air and a more reliable casting process.
- a relatively thin, approximately 0.015" (0.38 mm), refractory metal core element 10 is used to stabilize a tip region 12 of a ceramic core 14 during the casting process.
- the ceramic core 14 is positioned within a mold 80, only a portion of which has been shown.
- the ceramic core 14 may have the configuration of a laterally oriented passageway 15 to be formed in the airfoil tip region 34.
- the refractory metal core element 10 is printed out of the airfoil tip region 34 during casting and is located laterally of the ceramic core 14.
- the refractory metal core element 10 is positioned adjacent a side of the mold which forms the pressure side 40 of the airfoil portion 42.
- the refractory metal core element 10 is a metal piece which is much more rugged than typically brittle core print-outs. Thus, there is no manufacturing requirement for relatively large core print-out. A core print-out hole (not shown) may still be included if it is required for cooling purposes. In the present case, the core print-out hole can be made smaller than it previously could because it is not required to have as high of a strength. This configuration also allows for multiple ceramic core features to be stabilized by the same refractory metal core element. Furthermore, because this new tip design provides more stability and strength for the ceramic core 14 near the tip, the size of the trailing edge print-out of the tip flag cavity can be reduced, enabling lower cooling air flow out the tip flag exit.
- the refractory metal core element 10 may be formed from any suitable refractory material known in the art such as molybdenum or a molybdenum alloy.
- the refractory metal core element 10, as shown in FIGS. 2 , 4 and 5 may have a solid portion 46 and a plurality of spaced apart legs 48 depending downwardly from the solid portion 46.
- Each leg 48 preferably has a first leg portion 50, a base portion 52, and an angled portion 54 between the first portion 50 and the base portion 52.
- the base portion 52 of the legs may be joined together by a lower portion 53.
- the refractory metal core element 10 may be attached to the ceramic core 14 using any suitable means known in the art such as an adhesive or a mechanical fit connection.
- the refractory metal core element 10 and the ceramic core 14 are attached, inside the casting, the refractory metal core element can be used to control the location of both the refractory metal core and the ceramic core, relative to the external mold.
- the angled portion 54 may be omitted.
- the legs 48 can be arranged in any way that makes sense for the cooling design. Furthermore, the legs 48 only need to be connected at one end (inside or outside the casting), whichever makes sense for the cooling design and the casting process.
- the refractory metal core element 10 is printed out in such a way as to produce a row of aligned open slots 30 in the finished casting, along the pressure side edge 32 of the tip 34. Cooling air may be ejected from the slots 30 in whichever direction the slots 30 are oriented.
- the slots 30 may be oriented primarily radially outwards towards an outer circumference of the gaspath.
- the slots 30 may also be slightly angled towards the pressure side 40 of the turbine blade airfoil portion 42.
- the slots 30 may be purely radial or leaned in any combination of directions - forward/aft and/or towards pressure/suction side.
- the slots 30 may be in fluid communication with the passageway 15.
- the slots 30 are located in a recessed shelf 36 in the tip 34.
- the recessed shelf 36 may be a cast feature, or it may be machined into the finished casting in a later process.
- the cooling air When the cooling air exits the RMC defined tip slots 30, the cooling air immediately flows into a tip gap between the blade tip 34 and the blade outer air seal (BOAS)(not shown) due to the strong pressure gradient towards the suction side 60 of the airfoil portion 42. Injecting the cooling air into the tip gap significantly reduces the gaspath temperature in the tip gap downstream of the slots 30, resulting in lower heat load to the tip region of the blade. This is a similar effect to film cooling on the body of an airfoil.
- Conventional tip print-out holes provide some film cooling benefit on the tip surface, but they are significantly less efficient than this new design because the conventional tip print-out holes are so large that they can only be located at one or two locations along the mid-thickness of the tip.
- FIG. 3 shows a tip cooling design in accordance with the present invention which has only a single row of shaped cooling holes 70. The reduction of two rows of pressure side film cooling to one row is a benefit of the present invention, but it is not a necessary aspect of it.
- Tip blowing utilizes a row of cooling air jets or holes 30 along the pressure side edge 32 of the blade tip 34, which act to improve aerodynamic efficiency by reducing endwall losses associated with gaspath leakage across the tip gap.
- the cooling holes 70 are machined in the pressure side edge 32 after the blade and its airfoil portion have been cast.
- the cooling holes 70 may be machined using any suitable technique known in the art.
- the cooling holes 70 are in fluid communication with the passageway 15.
- the RMC-defined cooling slots 30 are situated along the recessed shelf 36 along the pressure side of the tip 34.
- the recessed shelf 36 will prevent the slots 30 from being unexpectedly closed during engine operation when the blade tip 34 rubs against the outer circumference of the gaspath.
- the recessed shelf 36 also allows for easier masking when applying abradable coating to the tip surface.
- the tip portion 34 of the airfoil portion 42 of the turbine engine blade is a cast structure and is formed at the same time as the remainder of the cast portions of the turbine engine blade.
- the mold 80 forming the tip region 34 of the airfoil portion 42 is illustrated in the drawings. It should be recognized that the mold 80 has a portion which is in the shape of the pressure side of the airfoil.
- the tip portion 34 may be formed by placing the ceramic core 14 into a mold 80. After the ceramic core 14, as well as any other needed ceramic or silica cores, has been positioned, the refractory metal core element 10 may be attached to the ceramic core 14 using any suitable means known in the art, such as an adhesive or pins.
- the mold 80 is created after the ceramic core 14 and the RMC 10 are assembled. This is preferably done by first assembling the ceramic core 14 and RMC 10, then injecting wax around the cores 10 and 14 using a wax die, so that the external surface of the wax is the same geometry as the external surface of finished casting. Then, a ceramic shell is applied to the external surface of the wax pattern. Then, the wax is melted out, leaving the ceramic core 14, RMC 10 and ceramic shell (not shown).
- the refractory metal core element 10 serves to stabilize the tip region of the ceramic core 14. Thereafter the blade with the airfoil portion may be cast using any suitable technique known in the art. After casting has been completed, the ceramic core 14 may be removed using any suitable technique known in the art to leave the passageway 15. Similarly, the refractory metal core element 10 is removed, thus leaving the slots 30. The RMC 10 may be leached out of the casting using any suitable chemical bath known in the art, very similar to how the ceramic cores are leached. Thereafter, a plurality of cooling holes 70 are machined into the tip region of the airfoil portion 42.
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- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Architecture (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
- Molds, Cores, And Manufacturing Methods Thereof (AREA)
Description
- The present invention relates to a turbine engine component, such as a turbine blade, having a plurality of as-cast blowing slots in a tip region.
- One of the typical failure modes for high pressure turbine (HPT) rotor airfoils (blades) is tip distress via oxidation and erosion. It is particularly challenging to design a cooling configuration for a tip region for a variety of reasons. First, it is very difficult to determine the external thermal boundary conditions near the tip due to the highly-three dimensional nature of the gaspath flow. Also, the tip region of a turbine blade is typically the thinnest portion of the airfoil, which makes it more difficult to package the desired cooling features. Furthermore, the tip region of a turbine blade is typically difficult to accurately produce with investment casting processes because the internal ceramic core is thin and weak near the tip. Further, it is cantilevered relatively far from the core-locating fixture at the blade root. Considering these points, it is desirable to have methods to create intricate cooling features near the tip capable of being targeted at specific regions of high heat load, while also allowing for greater control during the investment casting process.
- An existing HPT blade tip cooling design is shown in
FIG. 1 . A radially oriented cavity supplies cooling air to a leading edge impingement cooling scheme as well as a laterally-oriented cavity, known as a tip flag, that helps cool the tip before exiting the blade at the trailing edge near the tip.FIG. 1 also shows a midbody three-pass serpentine cooling arrangement and a trailing edge double-impingement system. - The tip of the core in
FIG. 1 includes an appendage that creates a recess blade tip known as a squealer pocket. That appendage is connected to the leading edge and tip flag core by means of two cylindrical connections ("print-outs") that form open holes in the finished casting ("print-out holes"). The core is fixed at the root of the blade during the casting process. The squealer pocket core is located laterally during the casting process, allowing the tip print outs to stabilize the tip region of the core. In order to prevent core breakage during the casting process, these tip print-outs should be as large as possible, especially considering that they are constructed from the brittle ceramic core material. One of the primary purposes of the squealer pocket is to allow for a shorter distance that the tip print-outs must span. However, it is desirable to have the tip print-out holes be smaller so that they do not flow an excessive amount of cooling air in the finished part, which results in inefficiency in the cooling design and, therefore, the turbine performance. -
EP-1445424-A2 describes an embedded microcircuit cooling design for a turbine blade tip. - In accordance with the present invention, there is provided a new tip cooling design that utilizes refractory metal core (RMC) technology in order to create a tip cooling scheme for a turbine engine component that is capable of more efficient use of cooling air and a more reliable casting process.
- In accordance with the present invention, there is provided a turbine engine component as defined by claim 1.
- There is also disclosed herein a process for forming an airfoil portion of a turbine engine component. The process comprises the steps of placing a ceramic core having a configuration of a passageway to be formed in the airfoil portion within a mold; attaching a refractory metal core element to the ceramic core to stabilize a tip region of the ceramic core during casting; and casting the airfoil portion.
- The process may further comprise locating the ceramic core relative to the mold with the refractory metal core element. The locating step may comprise providing a refractory metal core element having at least one leg. The refractory metal core element providing step may comprise providing a refractory metal core element having a plurality of legs. The process may further comprise removing the ceramic core so as to form the passageway and subsequently removing said refractory metal core element and thereby leaving at least one cooling slot in a tip region of the airfoil portion. The removing step may comprise leaving a plurality of cooling slots in said tip region. The process may further comprise machining a plurality of film cooling holes in the airfoil portion in the vicinity of the passageway formed by the ceramic core.
- There is also disclosed herein, in combination, a ceramic core for forming a passageway in a cast airfoil portion and means for stabilizing a tip region of the ceramic core. The stabilizing means comprises a refractory metal core element.
- The refractory metal core element may comprise a solid portion and a plurality of legs depending from the solid portion. Each leg may have an angled portion and a base portion and the base portion of the legs may be joined together by a lower portion.
- Still further, there is disclosed herein a refractory metal core element comprising a solid portion and a plurality of spaced apart legs depending from the solid portion. Each of the legs has a first portion adjacent the solid portion, a base portion, and an angled portion intermediate the first portion and the base portion so that the base portion is laterally offset from the solid portion. The base portions of the legs are preferably joined together by a lower portion.
- Other details of the RMC-defined tip blowing slots for turbine blade of the present invention, as well as other advantages attendant thereto, are set forth in the following detailed description and the accompanying drawings, wherein like reference numerals depict like elements.
-
-
FIG. 1 is a schematic representation of a cooling design used in a prior art turbine blade; -
FIG. 2 is an external view of a tip region of a casting; -
FIG. 3 is a schematic representation of a tip region of a cast airfoil portion of a turbine blade; -
FIG. 4 is a view of a refractory metal core element from the pressure side; and -
FIG. 5 is a view of the refractory metal core element from the trailing edge. - As noted before, a new tip cooling design for a turbine blade is proposed here that utilizes refractory metal core technology in order to help create a tip cooling scheme that is capable of more efficient use of cooling air and a more reliable casting process.
- Referring now to
FIG. 2 , a relatively thin, approximately 0.015" (0.38 mm), refractorymetal core element 10 is used to stabilize atip region 12 of aceramic core 14 during the casting process. Theceramic core 14 is positioned within amold 80, only a portion of which has been shown. Theceramic core 14 may have the configuration of a laterallyoriented passageway 15 to be formed in theairfoil tip region 34. The refractorymetal core element 10 is printed out of theairfoil tip region 34 during casting and is located laterally of theceramic core 14. Preferably, the refractorymetal core element 10 is positioned adjacent a side of the mold which forms thepressure side 40 of theairfoil portion 42. The refractorymetal core element 10 is a metal piece which is much more rugged than typically brittle core print-outs. Thus, there is no manufacturing requirement for relatively large core print-out. A core print-out hole (not shown) may still be included if it is required for cooling purposes. In the present case, the core print-out hole can be made smaller than it previously could because it is not required to have as high of a strength. This configuration also allows for multiple ceramic core features to be stabilized by the same refractory metal core element. Furthermore, because this new tip design provides more stability and strength for theceramic core 14 near the tip, the size of the trailing edge print-out of the tip flag cavity can be reduced, enabling lower cooling air flow out the tip flag exit. - The refractory
metal core element 10 may be formed from any suitable refractory material known in the art such as molybdenum or a molybdenum alloy. The refractorymetal core element 10, as shown inFIGS. 2 ,4 and 5 may have asolid portion 46 and a plurality of spacedapart legs 48 depending downwardly from thesolid portion 46. Eachleg 48 preferably has afirst leg portion 50, abase portion 52, and anangled portion 54 between thefirst portion 50 and thebase portion 52. Thebase portion 52 of the legs may be joined together by alower portion 53. The refractorymetal core element 10 may be attached to theceramic core 14 using any suitable means known in the art such as an adhesive or a mechanical fit connection. Because the refractorymetal core element 10 and theceramic core 14 are attached, inside the casting, the refractory metal core element can be used to control the location of both the refractory metal core and the ceramic core, relative to the external mold. In an alternative embodiment of the refractorymetal core element 10, theangled portion 54 may be omitted. Still further, thelegs 48 can be arranged in any way that makes sense for the cooling design. Furthermore, thelegs 48 only need to be connected at one end (inside or outside the casting), whichever makes sense for the cooling design and the casting process. - As shown in
FIG. 3 , the refractorymetal core element 10 is printed out in such a way as to produce a row of alignedopen slots 30 in the finished casting, along thepressure side edge 32 of thetip 34. Cooling air may be ejected from theslots 30 in whichever direction theslots 30 are oriented. As shown inFIG. 3 , theslots 30 may be oriented primarily radially outwards towards an outer circumference of the gaspath. Theslots 30 may also be slightly angled towards thepressure side 40 of the turbineblade airfoil portion 42. Theslots 30 may be purely radial or leaned in any combination of directions - forward/aft and/or towards pressure/suction side. Theslots 30 may be in fluid communication with thepassageway 15. As shown inFIG. 3 , theslots 30 are located in a recessedshelf 36 in thetip 34. The recessedshelf 36 may be a cast feature, or it may be machined into the finished casting in a later process. - When the cooling air exits the RMC defined
tip slots 30, the cooling air immediately flows into a tip gap between theblade tip 34 and the blade outer air seal (BOAS)(not shown) due to the strong pressure gradient towards thesuction side 60 of theairfoil portion 42. Injecting the cooling air into the tip gap significantly reduces the gaspath temperature in the tip gap downstream of theslots 30, resulting in lower heat load to the tip region of the blade. This is a similar effect to film cooling on the body of an airfoil. Conventional tip print-out holes provide some film cooling benefit on the tip surface, but they are significantly less efficient than this new design because the conventional tip print-out holes are so large that they can only be located at one or two locations along the mid-thickness of the tip. - Another cooling benefit of the RMC-defined
tip slots 30 is the substantial convective cooling of the pressure side region of thetip 34 due to the high-velocity cooling air flowing through thetip slots 30. This convective cooling is very effective at preventing oxidation and erosion along thepressure side edge 32 of thetip 34, which is a common location of tip distress. As a result of this increased convective cooling along thepressure side edge 32 of thetip 34, it is feasible to use fewer film cooling holes on the pressure side edge of the airfoil near the tip. In a prior art design, two rows of shaped cooling holes are provided along the pressure side near the tip. The purpose of these holes is to cool the tip region via film cooling and convective cooling.FIG. 3 shows a tip cooling design in accordance with the present invention which has only a single row of shaped cooling holes 70. The reduction of two rows of pressure side film cooling to one row is a benefit of the present invention, but it is not a necessary aspect of it. - The flexibility of the convective and film cooling aspects of the RMC-defined tip slots lends itself well to the challenge of designing a tip cooling configuration when the external boundary conditions are difficult to determine. Furthermore, the inherent strength of the refractory
metal core element 10 during the casting process allows for increased design flexibility in the tip region. As a result, this new tip cooling configuration allows for more efficient use of cooling air and more predictable casting yields, resulting in a more cost-effective product. - Another advantage of this tip cooling configuration is that it is complimentary to tip blowing technology for aerodynamic performance benefits. Tip blowing utilizes a row of cooling air jets or
holes 30 along thepressure side edge 32 of theblade tip 34, which act to improve aerodynamic efficiency by reducing endwall losses associated with gaspath leakage across the tip gap. The cooling holes 70 are machined in thepressure side edge 32 after the blade and its airfoil portion have been cast.
The cooling holes 70 may be machined using any suitable technique known in the art. The cooling holes 70 are in fluid communication with thepassageway 15. The RMC-definedcooling slots 30 are situated along the recessedshelf 36 along the pressure side of thetip 34. The recessedshelf 36 will prevent theslots 30 from being unexpectedly closed during engine operation when theblade tip 34 rubs against the outer circumference of the gaspath. The recessedshelf 36 also allows for easier masking when applying abradable coating to the tip surface. - The
tip portion 34 of theairfoil portion 42 of the turbine engine blade is a cast structure and is formed at the same time as the remainder of the cast portions of the turbine engine blade. For simplicity sake, only a portion of themold 80 forming thetip region 34 of theairfoil portion 42 is illustrated in the drawings. It should be recognized that themold 80 has a portion which is in the shape of the pressure side of the airfoil. - The
tip portion 34 may be formed by placing theceramic core 14 into amold 80. After theceramic core 14, as well as any other needed ceramic or silica cores, has been positioned, the refractorymetal core element 10 may be attached to theceramic core 14 using any suitable means known in the art, such as an adhesive or pins. Themold 80 is created after theceramic core 14 and theRMC 10 are assembled. This is preferably done by first assembling theceramic core 14 andRMC 10, then injecting wax around thecores ceramic core 14,RMC 10 and ceramic shell (not shown). As previously mentioned, the refractorymetal core element 10 serves to stabilize the tip region of theceramic core 14. Thereafter the blade with the airfoil portion may be cast using any suitable technique known in the art. After casting has been completed, theceramic core 14 may be removed using any suitable technique known in the art to leave thepassageway 15. Similarly, the refractorymetal core element 10 is removed, thus leaving theslots 30. TheRMC 10 may be leached out of the casting using any suitable chemical bath known in the art, very similar to how the ceramic cores are leached. Thereafter, a plurality of cooling holes 70 are machined into the tip region of theairfoil portion 42.
Claims (4)
- A turbine engine component having an airfoil portion (42) with a tip region (34), a recessed shelf (36) in said tip region (34), a plurality of slots (30) in said recessed shelf (36) through which a cooling fluid flows, said slots (30) being located along a pressure side (40) of said tip region (34), and further comprising a laterally oriented passageway (15) within said tip region (34), each of said slots (30) communicating with said passageway (15), characterised in that a plurality of cooling holes (70) are machined in said pressure side (40) and communicate with said passageway (15).
- The turbine engine component according to claim 1, further comprising said slots (30) being oriented primarily radially outwards and being angled towards said pressure side (40).
- The turbine engine component according to claim 1, wherein said passageway (15) comprises a laterally-oriented cavity.
- The turbine engine component according to claim 1, wherein the plurality of cooling holes (70) consists of a single row of shaped cooling holes (70).
Applications Claiming Priority (2)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US11/606,598 US20080131285A1 (en) | 2006-11-30 | 2006-11-30 | RMC-defined tip blowing slots for turbine blades |
EP07254584A EP1927414B1 (en) | 2006-11-30 | 2007-11-26 | RMC-Defined tip blowing slots for turbine blades |
Related Parent Applications (2)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
EP07254584A Division EP1927414B1 (en) | 2006-11-30 | 2007-11-26 | RMC-Defined tip blowing slots for turbine blades |
EP07254584.1 Division | 2007-11-26 |
Publications (2)
Publication Number | Publication Date |
---|---|
EP2246133A1 EP2246133A1 (en) | 2010-11-03 |
EP2246133B1 true EP2246133B1 (en) | 2014-07-09 |
Family
ID=39106243
Family Applications (2)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
EP10007545.6A Active EP2246133B1 (en) | 2006-11-30 | 2007-11-26 | RMC-defined tip blowing slots for turbine blades |
EP07254584A Active EP1927414B1 (en) | 2006-11-30 | 2007-11-26 | RMC-Defined tip blowing slots for turbine blades |
Family Applications After (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
EP07254584A Active EP1927414B1 (en) | 2006-11-30 | 2007-11-26 | RMC-Defined tip blowing slots for turbine blades |
Country Status (3)
Country | Link |
---|---|
US (1) | US20080131285A1 (en) |
EP (2) | EP2246133B1 (en) |
JP (1) | JP2008138675A (en) |
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-
2006
- 2006-11-30 US US11/606,598 patent/US20080131285A1/en not_active Abandoned
-
2007
- 2007-11-26 EP EP10007545.6A patent/EP2246133B1/en active Active
- 2007-11-26 EP EP07254584A patent/EP1927414B1/en active Active
- 2007-11-28 JP JP2007306709A patent/JP2008138675A/en active Pending
Cited By (1)
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US9579714B1 (en) | 2015-12-17 | 2017-02-28 | General Electric Company | Method and assembly for forming components having internal passages using a lattice structure |
Also Published As
Publication number | Publication date |
---|---|
EP1927414A3 (en) | 2008-07-30 |
EP1927414A2 (en) | 2008-06-04 |
EP1927414B1 (en) | 2013-01-23 |
JP2008138675A (en) | 2008-06-19 |
EP2246133A1 (en) | 2010-11-03 |
US20080131285A1 (en) | 2008-06-05 |
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