US20050265840A1 - Cooled rotor blade with leading edge impingement cooling - Google Patents
Cooled rotor blade with leading edge impingement cooling Download PDFInfo
- Publication number
- US20050265840A1 US20050265840A1 US10/855,076 US85507604A US2005265840A1 US 20050265840 A1 US20050265840 A1 US 20050265840A1 US 85507604 A US85507604 A US 85507604A US 2005265840 A1 US2005265840 A1 US 2005265840A1
- Authority
- US
- United States
- Prior art keywords
- rib
- radial passage
- crossover
- oblong
- rotor blade
- 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
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Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
- F01D5/187—Convection cooling
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- 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
- F05D2260/201—Heat transfer, e.g. cooling by impingement of a fluid
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- 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
- This invention applies to gas turbine rotor blades in general, and to cooled gas turbine rotor blades in particular.
- Turbine sections within an axial flow turbine engine include rotor assemblies that include a rotating disc and a number of rotor blades circumferentially disposed around the disk.
- Rotor blades include an airfoil portion for positioning within the gas path through the engine. Because the temperature within the gas path very often negatively affects the durability of the airfoil, it is known to cool an airfoil by passing cooling air through the airfoil. The cooled air helps decrease the temperature of the airfoil material and thereby increase its durability.
- Prior art cooled rotor blades very often utilize internal passage configurations that include a first radial passage extending contiguous with the leading edge, a second radial passage, and a rib disposed between and separating the passages.
- a plurality of crossover apertures is disposed within the rib, typically oriented perpendicular to the airfoil wall along the leading edge.
- a pressure difference across the rib causes a portion of the cooling air traveling within the second radial passage to pass through the crossover apertures and impinge on the leading edge wall.
- Prior art leading edge impingement configurations typically employed circular crossover apertures uniformly spaced along the rib.
- the cooling air impinging from each circular crossover aperture creates a region of relatively high heat transfer, albeit a small one.
- the circular crossover apertures create a line of discrete regions of high heat transfer separated by larger areas of relatively low heat transfer. The variations in heat transfer make the leading edge increase the possibility of undesirable fatigue, distress, oxidation, etc. within the leading edge wall.
- a rotor blade having a hollow airfoil and a root.
- the hollow airfoil has a cavity defined by a suction side wall, a pressure side wall, a leading edge, a trailing edge, a base, and a tip.
- An internal passage configuration is disposed within the cavity.
- the configuration includes a first radial passage, a second radial passage, and a rib disposed between and separating the first radial passage and second radial passage.
- a plurality of crossover apertures are disposed within the rib.
- a portion of the plurality of crossover apertures are oblong having a length extending through the rib, and a height and a width. The height of each oblong aperture is greater than the width.
- the oblong crossover apertures are aligned heightwise along the rib.
- the root includes a conduit that is operable to permit airflow through the root and into the first passage.
- One of the advantages of the present rotor blade and method is that airflow pressure losses within the airfoil are decreased relative to prior art airfoils having impingement cooling of which we are aware.
- FIG. 1 is a diagrammatic perspective view of the rotor assembly section.
- FIG. 2 is a diagrammatic sectional view of a rotor blade having an embodiment of the internal passage configuration.
- FIG. 3 is a diagrammatic sectional view of a portion of a rotor blade having an embodiment of the internal passage configuration.
- FIG. 4 is a diagrammatic partial view of a rib with oblong crossover apertures disposed therein.
- a rotor blade assembly 10 for a gas turbine engine having a disk 12 and a plurality of rotor blades 14 .
- the disk 12 includes a plurality of recesses 16 circumferentially disposed around the disk 12 and a rotational centerline 18 about which the disk 12 may rotate.
- Each blade 14 includes a root 20 , an airfoil 22 , a platform 24 , and a radial centerline 25 .
- the root 20 includes a geometry (e.g., a fir tree configuration) that mates with that of one of the recesses 16 within the disk 12 .
- the root 20 further includes conduits 26 through which cooling air may enter the root 20 and pass through into the airfoil 22 .
- the airfoil 22 includes a base 28 , a tip 30 , a leading edge 32 , a trailing edge 34 , a pressure side wall 36 (see FIG. 1 ), and a suction side wall 38 (see FIG. 1 ), and an internal passage configuration 40 .
- FIG. 2 diagrammatically illustrates an airfoil 22 sectioned between the leading edge 32 and the trailing edge 34 .
- the pressure side wall 36 and the suction side wall 38 extend between the base 28 and the tip 30 and meet at the leading edge 32 and the trailing edge 34 .
- the internal passage configuration includes a first conduit 42 , a second conduit 44 , and a third conduit 46 extending through the root 20 into the airfoil 22 . Fewer or more conduits may be used alternatively.
- the first conduit 42 is in fluid communication with a first radial passage 48 .
- a second radial passage 50 is disposed forward of the first radial passage 48 , contiguous with the leading edge 32 , and is connected to the first radial passage 48 by a plurality of crossover apertures 52 .
- the crossover apertures 52 are disposed in a rib 53 that extends between and separates the first radial passage 48 and the second radial passage 50 .
- the second radial passage 50 is connected to the exterior of the airfoil 22 by a plurality of cooling apertures 54 disposed along the leading edge 32 .
- the second radial passage 50 comprises one or more cavities.
- the second radial passage 50 may be in direct fluid communication with the first conduit 42 .
- the first radial passage 48 is connected to an axially extending passage 56 that extends to the trailing edge 34 of the airfoil 22 , adjacent the tip 30 of the airfoil 22 .
- a portion of the crossover apertures 52 disposed in the rib 53 are oblong, each having a length 70 , width 72 , and height 74 .
- substantially all of the crossover apertures 52 are oblong.
- the length 70 of each crossover aperture 52 extends through the rib 53 .
- the height 74 and width 72 are substantially perpendicular to each other and to the length 70 .
- the height 74 of each oblong crossover aperture 52 is greater than the width 72 . In a preferred embodiment, the height 74 is approximately twice the width 72 in magnitude.
- the oblong crossover apertures 52 are aligned heightwise along the rib 53 , such that the heights 74 of the oblong crossover apertures 52 are substantially collinear. In the embodiment shown in FIGS. 3 and 4 , the oblong crossover apertures 52 are shown as having a constant width 72 and circular ends. The oblong crossover apertures 52 are not limited to this embodiment.
- the rib 53 is separated from the interior surface of the leading edge wall 78 by a distance “L”.
- the oblong crossover apertures 52 may be described as having a hydraulic diameter “D”.
- the separation of the rib 53 from the leading edge wall 78 , and the size of the oblong crossover apertures 53 are such that the ratio of L/D is on average in the approximate range of 2.8 to 3.0. It is our experience that an L/D in this approximate range provides desirable impingement cooling.
- the first radial passage 48 includes a plurality of trip strips 58 attached to the interior surface of one or both of the pressure side wall 36 and the suction side wall 38 .
- the trip strips 58 are disposed within the passage 48 at an angle ⁇ that is skewed relative to the cooling airflow direction 60 within passage 48 ; i.e., at an angle between perpendicular and parallel to the airflow direction 60 .
- the trip strips 58 are oriented at angle of approximately 45° to the cooling airflow direction 60 .
- the orientation of each trip strip 58 within the passage 48 is such that the trip strip 58 converges toward the rib 53 containing the crossover apertures 52 , when viewed in the airflow direction 60 .
- Each of the trip strips 58 has an end disposed adjacent the rib 53 (i.e., a “rib end”). At least a portion of the trip strips 58 have a rib end radially located between a pair of crossover apertures 52 , preferably approximately midway between the pair of crossover apertures 52 .
- the second conduit 44 is in fluid communication with a serpentine passage 64 disposed immediately aft of the first and second radial passages 50 , 48 , in the mid-body region of the airfoil 22 .
- the serpentine passage 64 has an odd number of radial segments 66 , which number is greater than one; e.g., 3, 5, etc.
- the odd number of radial segments 66 ensures that the last radial segment in the serpentine 64 ends adjacent the axially extending passage 56 .
- Passage configurations other than the aforesaid serpentine passage 64 may be used within the mid-body region alternatively.
- the third conduit 46 is in fluid communication with one or more passages 68 disposed between the serpentine passage 64 and the trailing edge 34 of the airfoil 22 .
- the rotor blade airfoil 22 is disposed within the core gas path of the turbine engine.
- the airfoil 22 is subject to high temperature core gas passing by the airfoil 22 .
- Cooling air that is substantially lower in temperature than the core gas, is fed into the airfoil 22 through the conduits 42 , 44 , 46 disposed in the root 20 .
- Cooling air traveling through the first conduit 42 passes directly into the first radial passage 48 , and subsequently into the axially extending passage 56 adjacent the tip 30 of the airfoil 22 .
- a portion of the cooling air traveling within the first radial passage 48 encounters the trip strips 58 disposed within the passage 48 .
- the trip strips 58 converging toward the rib 53 direct the portion of cooling airflow toward the rib 53 .
- the position of the trip strips 58 relative to the crossover apertures 52 are such that the portion of cooling airflow directed toward the rib 53 is also directed toward the crossover apertures 52 .
- the portion of cooling airflow travels through the crossover apertures 52 and into the second radial passage 50 .
- the cooling air subsequently exits the second radial passage 50 via the cooling apertures 52 disposed in the leading edge 32 and impinges on the interior surface of the leading edge wall.
- prior art circular crossover apertures typically create a line of discrete regions of high heat transfer separated by larger areas of relatively low heat transfer.
- the oblong crossover apertures 52 of the present invention provide a more uniform radial heat transfer profile along the leading edge 32 that the aforesaid prior art.
- the regions of desirable relatively high heat transfer are larger, and the regions of undesirable relatively low heat transfer are smaller.
- the heat transfer within the regions of relatively low heat transfer appears to be increased by cooling air showering radially outward from the oblong crossover apertures 52 .
Abstract
A rotor blade is provided having a hollow airfoil and a root. The hollow airfoil has a cavity defined by a suction side wall, a pressure side wall, a leading edge, a trailing edge, a base, and a tip. An internal passage configuration is disposed within the cavity. The configuration includes a first radial passage, a second radial passage, and a rib disposed between and separating the first radial passage and second radial passage. A plurality of crossover apertures are disposed within the rib. A portion of the plurality of crossover apertures are oblong having a length extending through the rib, and a height and a width. The height of each oblong aperture is greater than the width. In some embodiments, the oblong crossover apertures are aligned heightwise along the rib. The root includes a conduit that is operable to permit airflow through the root and into the first radial passage.
Description
- 1. Technical Field
- This invention applies to gas turbine rotor blades in general, and to cooled gas turbine rotor blades in particular.
- 2. Background Information
- Turbine sections within an axial flow turbine engine include rotor assemblies that include a rotating disc and a number of rotor blades circumferentially disposed around the disk. Rotor blades include an airfoil portion for positioning within the gas path through the engine. Because the temperature within the gas path very often negatively affects the durability of the airfoil, it is known to cool an airfoil by passing cooling air through the airfoil. The cooled air helps decrease the temperature of the airfoil material and thereby increase its durability.
- Prior art cooled rotor blades very often utilize internal passage configurations that include a first radial passage extending contiguous with the leading edge, a second radial passage, and a rib disposed between and separating the passages. A plurality of crossover apertures is disposed within the rib, typically oriented perpendicular to the airfoil wall along the leading edge. A pressure difference across the rib causes a portion of the cooling air traveling within the second radial passage to pass through the crossover apertures and impinge on the leading edge wall.
- Prior art leading edge impingement configurations typically employed circular crossover apertures uniformly spaced along the rib. The cooling air impinging from each circular crossover aperture creates a region of relatively high heat transfer, albeit a small one. Collectively, the circular crossover apertures create a line of discrete regions of high heat transfer separated by larger areas of relatively low heat transfer. The variations in heat transfer make the leading edge increase the possibility of undesirable fatigue, distress, oxidation, etc. within the leading edge wall.
- What is needed is an airfoil having improved impingement cooling that increases the uniformity of impingement cooling, particularly along the leading edge of the blade.
- According to the present invention, a rotor blade is provided having a hollow airfoil and a root. The hollow airfoil has a cavity defined by a suction side wall, a pressure side wall, a leading edge, a trailing edge, a base, and a tip. An internal passage configuration is disposed within the cavity. The configuration includes a first radial passage, a second radial passage, and a rib disposed between and separating the first radial passage and second radial passage. A plurality of crossover apertures are disposed within the rib. A portion of the plurality of crossover apertures are oblong having a length extending through the rib, and a height and a width. The height of each oblong aperture is greater than the width. In some embodiments, the oblong crossover apertures are aligned heightwise along the rib. The root includes a conduit that is operable to permit airflow through the root and into the first passage.
- One of the advantages of the present rotor blade and method is that airflow pressure losses within the airfoil are decreased relative to prior art airfoils having impingement cooling of which we are aware.
- These and other objects, features and advantages of the present invention will become apparent in light of the detailed description of the best mode embodiment thereof, as illustrated in the accompanying drawings.
-
FIG. 1 is a diagrammatic perspective view of the rotor assembly section. -
FIG. 2 is a diagrammatic sectional view of a rotor blade having an embodiment of the internal passage configuration. -
FIG. 3 is a diagrammatic sectional view of a portion of a rotor blade having an embodiment of the internal passage configuration. -
FIG. 4 is a diagrammatic partial view of a rib with oblong crossover apertures disposed therein. - Referring to
FIG. 1 , arotor blade assembly 10 for a gas turbine engine is provided having adisk 12 and a plurality ofrotor blades 14. Thedisk 12 includes a plurality ofrecesses 16 circumferentially disposed around thedisk 12 and arotational centerline 18 about which thedisk 12 may rotate. Eachblade 14 includes aroot 20, anairfoil 22, aplatform 24, and aradial centerline 25. Theroot 20 includes a geometry (e.g., a fir tree configuration) that mates with that of one of therecesses 16 within thedisk 12. As can be seen inFIG. 2 , theroot 20 further includesconduits 26 through which cooling air may enter theroot 20 and pass through into theairfoil 22. - Referring to
FIGS. 2-4 , theairfoil 22 includes abase 28, atip 30, a leadingedge 32, atrailing edge 34, a pressure side wall 36 (seeFIG. 1 ), and a suction side wall 38 (seeFIG. 1 ), and aninternal passage configuration 40.FIG. 2 diagrammatically illustrates anairfoil 22 sectioned between the leadingedge 32 and thetrailing edge 34. Thepressure side wall 36 and thesuction side wall 38 extend between thebase 28 and thetip 30 and meet at the leadingedge 32 and thetrailing edge 34. - The internal passage configuration includes a
first conduit 42, asecond conduit 44, and athird conduit 46 extending through theroot 20 into theairfoil 22. Fewer or more conduits may be used alternatively. Thefirst conduit 42 is in fluid communication with a firstradial passage 48. A secondradial passage 50 is disposed forward of the firstradial passage 48, contiguous with the leadingedge 32, and is connected to the firstradial passage 48 by a plurality ofcrossover apertures 52. Thecrossover apertures 52 are disposed in arib 53 that extends between and separates the firstradial passage 48 and the secondradial passage 50. The secondradial passage 50 is connected to the exterior of theairfoil 22 by a plurality ofcooling apertures 54 disposed along the leadingedge 32. In some embodiments, the secondradial passage 50 comprises one or more cavities. In other embodiments, the secondradial passage 50 may be in direct fluid communication with thefirst conduit 42. At the outer radial end of the first radial passage 48 (i.e., the end of the firstradial passage 48 opposite the first conduit 42), the firstradial passage 48 is connected to an axially extendingpassage 56 that extends to thetrailing edge 34 of theairfoil 22, adjacent thetip 30 of theairfoil 22. - A portion of the
crossover apertures 52 disposed in therib 53 are oblong, each having alength 70,width 72, andheight 74. In a preferred embodiment, substantially all of thecrossover apertures 52 are oblong. Thelength 70 of eachcrossover aperture 52 extends through therib 53. Theheight 74 andwidth 72 are substantially perpendicular to each other and to thelength 70. Theheight 74 of eachoblong crossover aperture 52 is greater than thewidth 72. In a preferred embodiment, theheight 74 is approximately twice thewidth 72 in magnitude. Theoblong crossover apertures 52 are aligned heightwise along therib 53, such that theheights 74 of theoblong crossover apertures 52 are substantially collinear. In the embodiment shown inFIGS. 3 and 4 , theoblong crossover apertures 52 are shown as having aconstant width 72 and circular ends. Theoblong crossover apertures 52 are not limited to this embodiment. - The
rib 53 is separated from the interior surface of the leadingedge wall 78 by a distance “L”. Theoblong crossover apertures 52 may be described as having a hydraulic diameter “D”. In a preferred embodiment, the separation of therib 53 from the leadingedge wall 78, and the size of theoblong crossover apertures 53 are such that the ratio of L/D is on average in the approximate range of 2.8 to 3.0. It is our experience that an L/D in this approximate range provides desirable impingement cooling. - The first
radial passage 48 includes a plurality of trip strips 58 attached to the interior surface of one or both of thepressure side wall 36 and thesuction side wall 38. The trip strips 58 are disposed within thepassage 48 at an angle α that is skewed relative to thecooling airflow direction 60 withinpassage 48; i.e., at an angle between perpendicular and parallel to theairflow direction 60. Preferably, the trip strips 58 are oriented at angle of approximately 45° to thecooling airflow direction 60. The orientation of eachtrip strip 58 within thepassage 48 is such that thetrip strip 58 converges toward therib 53 containing thecrossover apertures 52, when viewed in theairflow direction 60. Each of the trip strips 58 has an end disposed adjacent the rib 53 (i.e., a “rib end”). At least a portion of the trip strips 58 have a rib end radially located between a pair ofcrossover apertures 52, preferably approximately midway between the pair ofcrossover apertures 52. - Referring to
FIG. 2 , thesecond conduit 44 is in fluid communication with aserpentine passage 64 disposed immediately aft of the first and secondradial passages airfoil 22. Theserpentine passage 64 has an odd number of radial segments 66, which number is greater than one; e.g., 3, 5, etc. The odd number of radial segments 66 ensures that the last radial segment in the serpentine 64 ends adjacent theaxially extending passage 56. Passage configurations other than theaforesaid serpentine passage 64 may be used within the mid-body region alternatively. - The
third conduit 46 is in fluid communication with one ormore passages 68 disposed between theserpentine passage 64 and the trailingedge 34 of theairfoil 22. - In the operation of the invention, the
rotor blade airfoil 22 is disposed within the core gas path of the turbine engine. Theairfoil 22 is subject to high temperature core gas passing by theairfoil 22. Cooling air, that is substantially lower in temperature than the core gas, is fed into theairfoil 22 through theconduits root 20. p Cooling air traveling through thefirst conduit 42 passes directly into the firstradial passage 48, and subsequently into theaxially extending passage 56 adjacent thetip 30 of theairfoil 22. A portion of the cooling air traveling within the firstradial passage 48 encounters the trip strips 58 disposed within thepassage 48. The trip strips 58 converging toward therib 53 direct the portion of cooling airflow toward therib 53. The position of the trip strips 58 relative to thecrossover apertures 52 are such that the portion of cooling airflow directed toward therib 53 is also directed toward thecrossover apertures 52. The portion of cooling airflow travels through thecrossover apertures 52 and into the secondradial passage 50. The cooling air subsequently exits the secondradial passage 50 via thecooling apertures 52 disposed in the leadingedge 32 and impinges on the interior surface of the leading edge wall. - As stated above, prior art circular crossover apertures typically create a line of discrete regions of high heat transfer separated by larger areas of relatively low heat transfer. The
oblong crossover apertures 52 of the present invention provide a more uniform radial heat transfer profile along the leadingedge 32 that the aforesaid prior art. The regions of desirable relatively high heat transfer are larger, and the regions of undesirable relatively low heat transfer are smaller. In addition, the heat transfer within the regions of relatively low heat transfer appears to be increased by cooling air showering radially outward from theoblong crossover apertures 52. - Although this invention has been shown and described with respect to the detailed embodiments thereof, it will be understood by those skilled in the art that various changes in form and detail thereof may be made without departing from the spirit and the scope of the invention.
Claims (9)
1. A rotor blade, comprising:
a hollow airfoil having a cavity defined by a suction side wall, a pressure side wall, a leading edge, a trailing edge, a base, and a tip;
an internal passage configuration disposed within the cavity, which configuration includes a first radial passage, a second radial passage, a rib disposed between and separating the first radial passage and second radial passage, and a plurality of crossover apertures disposed within the rib, wherein a portion of the plurality of crossover apertures are oblong having a length extending through the rib, and a height and a width, and wherein the height of each oblong crossover aperture is greater than the width;
a root having a conduit that is operable to permit airflow through the root and into the first radial passage.
2. The rotor blade of claim 1 , wherein substantially all of the crossover apertures are oblong.
3. The rotor blade of claim 2 , wherein the height of each crossover aperture is approximately twice the magnitude of the width of that crossover aperture.
4. The rotor blade of claim 1 , wherein the oblong crossover apertures are aligned heightwise along the rib.
5. The rotor blade of claim 4 , wherein the second radial passage is contiguous with the leading edge.
6. A rotor blade, comprising:
a hollow airfoil having a cavity defined by a suction side wall, a pressure side wall, a leading edge, a trailing edge, a base, and a tip;
an internal passage configuration disposed within the cavity, which configuration includes a first radial passage, a second radial passage, a rib disposed between and separating the first radial passage and second radial passage, and a plurality of crossover apertures disposed within the rib, wherein a portion of the plurality of crossover apertures are oblong having a length extending through the rib, and a height and a width, and wherein the height of each oblong crossover aperture is greater than the width, and wherein the rib is separated from the leading edge by a distance “L”, and the oblong crossover apertures have a hydraulic diameter “D”, and the ratio of L/D is on average in the approximate range of 2.8 to 3.0; and
a root having a conduit that is operable to permit airflow through the root and into the first radial passage.
7. The rotor blade of claim 6 , wherein substantially all of the crossover apertures are oblong.
8. The rotor blade of claim 7 , wherein the height of each crossover aperture is approximately twice the magnitude of the width of that crossover aperture.
9. The rotor blade of claim 6 , wherein the oblong crossover apertures are aligned heightwise along the rib.
Priority Applications (4)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US10/855,076 US20050265840A1 (en) | 2004-05-27 | 2004-05-27 | Cooled rotor blade with leading edge impingement cooling |
JP2005154978A JP2005337257A (en) | 2004-05-27 | 2005-05-27 | Rotor blade |
EP05253261A EP1605138B1 (en) | 2004-05-27 | 2005-05-27 | Cooled rotor blade with leading edge impingement cooling |
DE602005022018T DE602005022018D1 (en) | 2004-05-27 | 2005-05-27 | Cooled rotor blade with impingement cooling in the area of the leading edge |
Applications Claiming Priority (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US10/855,076 US20050265840A1 (en) | 2004-05-27 | 2004-05-27 | Cooled rotor blade with leading edge impingement cooling |
Publications (1)
Publication Number | Publication Date |
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US20050265840A1 true US20050265840A1 (en) | 2005-12-01 |
Family
ID=34941473
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
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US10/855,076 Abandoned US20050265840A1 (en) | 2004-05-27 | 2004-05-27 | Cooled rotor blade with leading edge impingement cooling |
Country Status (4)
Country | Link |
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US (1) | US20050265840A1 (en) |
EP (1) | EP1605138B1 (en) |
JP (1) | JP2005337257A (en) |
DE (1) | DE602005022018D1 (en) |
Cited By (9)
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US20060083614A1 (en) * | 2004-10-18 | 2006-04-20 | United Technologies Corporation | Airfoil with large fillet and micro-circuit cooling |
US20130280080A1 (en) * | 2012-04-23 | 2013-10-24 | Jeffrey R. Levine | Gas turbine engine airfoil with dirt purge feature and core for making same |
US20140238028A1 (en) * | 2011-11-08 | 2014-08-28 | Ihi Corporation | Impingement cooling mechanism, turbine blade, and combustor |
US8840370B2 (en) | 2011-11-04 | 2014-09-23 | General Electric Company | Bucket assembly for turbine system |
US20160024938A1 (en) * | 2014-07-25 | 2016-01-28 | United Technologies Corporation | Airfoil cooling apparatus |
US20160115796A1 (en) * | 2013-05-20 | 2016-04-28 | Kawasaki Jukogyo Kabushiki Kaisha | Turbine blade cooling structure |
US9932836B2 (en) | 2012-03-22 | 2018-04-03 | Ansaldo Energia Ip Uk Limited | Turbine blade |
US20180298763A1 (en) * | 2014-11-11 | 2018-10-18 | Siemens Aktiengesellschaft | Turbine blade with axial tip cooling circuit |
EP3597859B1 (en) * | 2018-07-13 | 2023-08-30 | Honeywell International Inc. | Turbine blade with dust tolerant cooling system |
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FR2918105B1 (en) * | 2007-06-27 | 2013-12-27 | Snecma | TURBOMACHINE COOLED AUBE COMPRISING VARIABLE IMPACT REMOTE COOLING HOLES. |
US9151173B2 (en) * | 2011-12-15 | 2015-10-06 | General Electric Company | Use of multi-faceted impingement openings for increasing heat transfer characteristics on gas turbine components |
KR101906701B1 (en) * | 2017-01-03 | 2018-10-10 | 두산중공업 주식회사 | Gas turbine blade |
US10895168B2 (en) * | 2019-05-30 | 2021-01-19 | Solar Turbines Incorporated | Turbine blade with serpentine channels |
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US5688104A (en) * | 1993-11-24 | 1997-11-18 | United Technologies Corporation | Airfoil having expanded wall portions to accommodate film cooling holes |
KR20000052372A (en) * | 1999-01-25 | 2000-08-25 | 제이 엘. 차스킨, 버나드 스나이더, 아더엠. 킹 | Gas turbine bucket cooling passage connectors |
US6174134B1 (en) * | 1999-03-05 | 2001-01-16 | General Electric Company | Multiple impingement airfoil cooling |
US6290463B1 (en) * | 1999-09-30 | 2001-09-18 | General Electric Company | Slotted impingement cooling of airfoil leading edge |
EP1213442B1 (en) * | 2000-12-05 | 2009-03-11 | United Technologies Corporation | Rotor blade |
DE10332563A1 (en) * | 2003-07-11 | 2005-01-27 | Rolls-Royce Deutschland Ltd & Co Kg | Turbine blade with impingement cooling |
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-
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- 2005-05-27 DE DE602005022018T patent/DE602005022018D1/en active Active
- 2005-05-27 EP EP05253261A patent/EP1605138B1/en active Active
- 2005-05-27 JP JP2005154978A patent/JP2005337257A/en active Pending
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US3767322A (en) * | 1971-07-30 | 1973-10-23 | Westinghouse Electric Corp | Internal cooling for turbine vanes |
US5387086A (en) * | 1993-07-19 | 1995-02-07 | General Electric Company | Gas turbine blade with improved cooling |
US6705836B2 (en) * | 2001-08-28 | 2004-03-16 | Snecma Moteurs | Gas turbine blade cooling circuits |
Cited By (14)
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US20060083614A1 (en) * | 2004-10-18 | 2006-04-20 | United Technologies Corporation | Airfoil with large fillet and micro-circuit cooling |
US7217094B2 (en) * | 2004-10-18 | 2007-05-15 | United Technologies Corporation | Airfoil with large fillet and micro-circuit cooling |
US8840370B2 (en) | 2011-11-04 | 2014-09-23 | General Electric Company | Bucket assembly for turbine system |
US20140238028A1 (en) * | 2011-11-08 | 2014-08-28 | Ihi Corporation | Impingement cooling mechanism, turbine blade, and combustor |
US9932836B2 (en) | 2012-03-22 | 2018-04-03 | Ansaldo Energia Ip Uk Limited | Turbine blade |
US9279331B2 (en) * | 2012-04-23 | 2016-03-08 | United Technologies Corporation | Gas turbine engine airfoil with dirt purge feature and core for making same |
US20130280080A1 (en) * | 2012-04-23 | 2013-10-24 | Jeffrey R. Levine | Gas turbine engine airfoil with dirt purge feature and core for making same |
US9938837B2 (en) | 2012-04-23 | 2018-04-10 | United Technologies Corporation | Gas turbine engine airfoil trailing edge passage and core for making same |
US20160115796A1 (en) * | 2013-05-20 | 2016-04-28 | Kawasaki Jukogyo Kabushiki Kaisha | Turbine blade cooling structure |
US10018053B2 (en) * | 2013-05-20 | 2018-07-10 | Kawasaki Jukogyo Kabushiki Kaisha | Turbine blade cooling structure |
US20160024938A1 (en) * | 2014-07-25 | 2016-01-28 | United Technologies Corporation | Airfoil cooling apparatus |
US10012090B2 (en) * | 2014-07-25 | 2018-07-03 | United Technologies Corporation | Airfoil cooling apparatus |
US20180298763A1 (en) * | 2014-11-11 | 2018-10-18 | Siemens Aktiengesellschaft | Turbine blade with axial tip cooling circuit |
EP3597859B1 (en) * | 2018-07-13 | 2023-08-30 | Honeywell International Inc. | Turbine blade with dust tolerant cooling system |
Also Published As
Publication number | Publication date |
---|---|
JP2005337257A (en) | 2005-12-08 |
EP1605138A3 (en) | 2007-10-03 |
EP1605138A2 (en) | 2005-12-14 |
DE602005022018D1 (en) | 2010-08-12 |
EP1605138B1 (en) | 2010-06-30 |
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