US20090208343A1 - Serpentine microcircuits for hot gas migration - Google Patents
Serpentine microcircuits for hot gas migration Download PDFInfo
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- US20090208343A1 US20090208343A1 US11/494,831 US49483106A US2009208343A1 US 20090208343 A1 US20090208343 A1 US 20090208343A1 US 49483106 A US49483106 A US 49483106A US 2009208343 A1 US2009208343 A1 US 2009208343A1
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- United States
- Prior art keywords
- cooling
- pressure side
- turbine engine
- engine component
- circuit
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- 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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Classifications
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
- F01D5/187—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
- F05D2260/00—Function
- F05D2260/20—Heat transfer, e.g. cooling
- F05D2260/221—Improvement of heat transfer
Definitions
- the present invention relates to a turbine engine component having an improved scheme for cooling an airfoil portion.
- the overall cooling effectiveness is a measure used to determine the cooling characteristics of a particular design.
- the ideal non-achievable goal is unity, which implies that the metal temperature is the same as the coolant temperature inside an airfoil.
- the opposite can also occur when the cooling effectiveness is zero implying that the metal temperature is the same as the gas temperature. In that case, the blade material will certainly melt and burn away.
- existing cooling technology allows the cooling effectiveness to be between 0.5 and 0.6. More advanced technology such as supercooling should be between 0.6 and 0.7. Microcircuit cooling as the most advanced cooling technology in existence today can be made to produce cooling effectiveness higher than 0.7.
- FIG. 1 shows a durability map of cooling effectiveness (x-axis) vs. the film effectiveness (y-axis) for different lines of convective efficiency. Placed in the map is a point 10 related to a new advanced serpentine microcircuit shown in FIGS. 2 a - 2 c.
- This serpentine microcircuit includes a pressure side serpentine circuit 20 and a suction side serpentine circuit 22 embedded in the airfoil walls 24 and 26 .
- the Table I below provides the operational parameters used to plot the design point in the durability map.
- FIG. 3 illustrates the cooling flow distribution for a turbine blade with the serpentine microcircuits of FIGS. 2 a - 2 c embedded in the airfoils walls.
- FIGS. 4A and 4B There are however field problems that can be addressed efficiently with peripheral microcircuit designs.
- FIG. 4A the streamlines of the gas path close to the external surface of the airfoil illustrate four different regions in which the gas flow changes direction or migration: a tip region, two mid-section regions, and a root region. In between the tip and the upper mid region, the flow transitions through a pseudo stagnation point(s). The momentum of the external gas seems to decelerate in such a way as to impose a local thermal load to the part. This manifests itself by regions where the propensity for erosion and oxidation increase in the airfoil surface. The superposition of FIG.
- 4B illustrates the local coincidence between the pseudo-stagnation region and the blade distress in the part surface.
- the upper and lower region also converge onto one another, but even though the space between streamlines decreases, the flow seems to accelerate and there is no pseudo-stagnation regions.
- a mild manifestation of the same tip-to-mid phenomena seems to initiate in the transition region between the mid-to-root regions. It is therefore necessary to tailor the peripheral microcircuit in such a manner as to address these local high thermal load regions.
- a turbine engine component is provided with improved cooling.
- the turbine engine component broadly comprises an airfoil portion having a pressure side and a suction side.
- the turbine engine component further has a first cooling circuit within the pressure side for cooling the pressure side of the airfoil portion and a second cooling circuit within the suction side for cooling the suction side of the airfoil portion and for cooperating with means for creating a cooling film over the pressure side.
- FIG. 1 is a graph showing cooling effectiveness versus film effectiveness for a turbine engine component
- FIG. 2A shows an airfoil portion of a turbine engine component having a pressure side cooling microcircuit embedded in the pressure side wall and a suction side cooling microcircuit embedded in the suction side wall;
- FIG. 2B is a schematic representation of a pressure side cooling microcircuit used in the airfoil portion of FIG. 2A ;
- FIG. 2C is a schematic representation of a suction side cooling microcircuit used in the airfoil portion of FIG. 2A ;
- FIG. 3 illustrates the cooling flow distribution for a turbine engine component with serpentine microcircuits embedded in the airfoil walls
- FIG. 4A is a schematic representation illustrating the pressure side distress on an airfoil surface
- FIG. 4B is a schematic representation of the local coincidence between the pseudo-stagnation region and the blade distress
- FIG. 5 is a schematic representation of a peripheral pressure side cooling circuit
- FIG. 6 is a schematic representation of a peripheral suction side cooling circuit
- FIG. 7 is a schematic representation of main body internal cooling circuits.
- the two peripheral cooling arrangements include a peripheral pressure side microcircuit 100 which may be incorporated or embedded within the wall forming the pressure side of an airfoil portion 104 and a suction side microcircuit 120 which may be incorporated or embedded within the wall forming the suction side of the airfoil portion 104 .
- the pressure side peripheral microcircuit 100 is shown.
- the first leg 102 has an inlet 103 which receives cooling fluid from a source (not shown).
- the leg 102 provides a flow of cooling fluid which quenches the hot spot in the tip-to-mid region of the airfoil portion 104 shown in FIG. 4B .
- the cooling fluid within the leg 102 proceeds around a 180 degree bend 106 which is supplemented with a plurality of film holes 108 , preferably three film holes.
- the film holes 108 ensure flow acceleration through the bend 106 to a second downstream leg 110 which ends below the platform 112 of the turbine engine component 90 in an exit 164 . Cooling fluid from the leg 110 is fed into an internal trailing edge circuit 114 to be discussed hereinafter via the exit 164 where it is used to further cool the airfoil portion 104 .
- the circuit 120 has a first leg 122 which communicates with a source (not shown) of cooling fluid. In the first leg 122 , the cooling flow convects heat away from the suction side. Since the circuit 120 has no film holes, effective cooling may not be done past the external gage point of the airfoil portion 104 where any film cooling would provide high aerodynamic penalties due to mixing. (PLEASE CHECK THIS TO SEE IF IT MAKES SENSE) Thus, the circuit 120 is used to feed cooling fluid to a leading edge microcircuit 124 which wraps around the leading edge 126 of the airfoil portion 104 .
- the circuit 120 feeds or supplies cooling fluid to the leading edge wrap around circuit 124 through a plurality of wall cross over holes 128 .
- the circuit 120 has a bend 130 and a second leg 132 .
- the holes 128 are preferably located in the vicinity of the bend 130 and the second leg 132 .
- the second leg 132 may also communicate with the wrap around circuit 124 via a passageway 134 .
- the main body internal cooling circuits which include a leading edge internal cooling circuit 150 and the trailing edge internal cooling circuit 114 .
- the leading edge internal cooling circuit 150 communicates with a source (not shown) of cooling fluid, such as engine bleed air, via an inlet 151 and has one or more film cooling holes 152 adjacent the tip 154 of the airfoil portion 104 to provide tip cooling.
- the circuit 150 also has a plurality of cross-over holes 156 for supplying cooling fluid to the leading edge microcircuit 124 .
- the trailing edge internal circuit 114 also communicates with a source (not shown) of cooling fluid, such as engine bleed air, via an inlet 157 and has one or more film cooling holes 158 adjacent the tip 154 to provide tip cooling.
- the circuit 114 also has a plurality of cross-over holes 160 for communicating with a trailing edge cooling circuit 162 for cooling the trailing edge of the airfoil portion 104 .
- the tailing edge internal circuit 114 also receives cooling fluid from the peripheral pressure side microcircuit 100 via the exit 164 .
- Each of the leading edge internal circuit 150 and the trailing edge internal circuit 114 may be provided with a plurality of film cooling holes 170 and 172 respectively to form cooling films over the pressure and suction sides of the airfoil portion 104 .
- the airfoil portion of a turbine engine component may be very effectively convectively cooled.
- the cooling flow is returned to the trailing edge internal circuit for further cooling of the airfoil.
- the suction side circuit the leading edge of the airfoil is cooled first before discharging in pressure side film. This effective use of coolant allows for positive effects on cycle thermodynamic efficiency, turbine efficiency, rotor inlet temperature impacts, and specific fuel consumption.
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- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
Description
- (1) Field of the Invention
- The present invention relates to a turbine engine component having an improved scheme for cooling an airfoil portion.
- (2) Prior Art
- The overall cooling effectiveness is a measure used to determine the cooling characteristics of a particular design. The ideal non-achievable goal is unity, which implies that the metal temperature is the same as the coolant temperature inside an airfoil. The opposite can also occur when the cooling effectiveness is zero implying that the metal temperature is the same as the gas temperature. In that case, the blade material will certainly melt and burn away. In general, existing cooling technology allows the cooling effectiveness to be between 0.5 and 0.6. More advanced technology such as supercooling should be between 0.6 and 0.7. Microcircuit cooling as the most advanced cooling technology in existence today can be made to produce cooling effectiveness higher than 0.7.
-
FIG. 1 shows a durability map of cooling effectiveness (x-axis) vs. the film effectiveness (y-axis) for different lines of convective efficiency. Placed in the map is apoint 10 related to a new advanced serpentine microcircuit shown inFIGS. 2 a-2 c. This serpentine microcircuit includes a pressureside serpentine circuit 20 and a suctionside serpentine circuit 22 embedded in theairfoil walls - The Table I below provides the operational parameters used to plot the design point in the durability map.
-
TABLE I Operational Parameters for serpentine microcircuit beta 2.898 Tg 2581 [F.] Tc 1365 [F.] Tm 2050 [F.] Tm_bulk 1709 [F.] Phi_loc 0.437 Phi_bulk 0.717 Tco 1640 [F.] Tci 1090 [F.] eta_c_loc 0.573 eta_f 0.296 Total Cooling Flow 3.503% WAE 10.8 Legend for Table I Beta = heat load Phi_loc = local cooling effectiveness Phi_bulk = bulk cooling effectiveness Eta_c_loc = local cooling efficiency Eta_f = film effectiveness Tg = gas temperature Tc = coolant temperature Tm = metal temperature Tm_bulk = bulk metal temperature Tco = exit coolant temperature Tci = inlet coolant temperature WAE = compressor engine flow, pps - It should be noted that the overall cooling effectiveness from the table is 0.717 for a film effectiveness of 0.296 and a convective efficiency (or ability to pick-up heat) of 0.573. Also note that the corresponding cooling flow for a turbine blade having this cooling microcircuit is 3.5% engine flow.
FIG. 3 illustrates the cooling flow distribution for a turbine blade with the serpentine microcircuits ofFIGS. 2 a-2 c embedded in the airfoils walls. - There are however field problems that can be addressed efficiently with peripheral microcircuit designs. One such field problem is illustrated in
FIGS. 4A and 4B . InFIG. 4A , the streamlines of the gas path close to the external surface of the airfoil illustrate four different regions in which the gas flow changes direction or migration: a tip region, two mid-section regions, and a root region. In between the tip and the upper mid region, the flow transitions through a pseudo stagnation point(s). The momentum of the external gas seems to decelerate in such a way as to impose a local thermal load to the part. This manifests itself by regions where the propensity for erosion and oxidation increase in the airfoil surface. The superposition ofFIG. 4B illustrates the local coincidence between the pseudo-stagnation region and the blade distress in the part surface. In the mid region, the upper and lower region also converge onto one another, but even though the space between streamlines decreases, the flow seems to accelerate and there is no pseudo-stagnation regions. A mild manifestation of the same tip-to-mid phenomena seems to initiate in the transition region between the mid-to-root regions. It is therefore necessary to tailor the peripheral microcircuit in such a manner as to address these local high thermal load regions. - In accordance with the present invention, a turbine engine component is provided with improved cooling. The turbine engine component broadly comprises an airfoil portion having a pressure side and a suction side. The turbine engine component further has a first cooling circuit within the pressure side for cooling the pressure side of the airfoil portion and a second cooling circuit within the suction side for cooling the suction side of the airfoil portion and for cooperating with means for creating a cooling film over the pressure side.
- Other details of the serpentine microcircuits for hot gas migration of the present invention, as well as other objects and 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 graph showing cooling effectiveness versus film effectiveness for a turbine engine component; -
FIG. 2A shows an airfoil portion of a turbine engine component having a pressure side cooling microcircuit embedded in the pressure side wall and a suction side cooling microcircuit embedded in the suction side wall; -
FIG. 2B is a schematic representation of a pressure side cooling microcircuit used in the airfoil portion ofFIG. 2A ; -
FIG. 2C is a schematic representation of a suction side cooling microcircuit used in the airfoil portion ofFIG. 2A ; -
FIG. 3 illustrates the cooling flow distribution for a turbine engine component with serpentine microcircuits embedded in the airfoil walls; -
FIG. 4A is a schematic representation illustrating the pressure side distress on an airfoil surface; -
FIG. 4B is a schematic representation of the local coincidence between the pseudo-stagnation region and the blade distress; -
FIG. 5 is a schematic representation of a peripheral pressure side cooling circuit; -
FIG. 6 is a schematic representation of a peripheral suction side cooling circuit; and -
FIG. 7 is a schematic representation of main body internal cooling circuits. - Referring now to
FIGS. 5 and 6 , there are depicted two peripheral cooling arrangements which may be used to address local increases in the airfoil thermal load of aturbine engine component 90 such as a turbine blade. The two peripheral cooling arrangements include a peripheralpressure side microcircuit 100 which may be incorporated or embedded within the wall forming the pressure side of anairfoil portion 104 and asuction side microcircuit 120 which may be incorporated or embedded within the wall forming the suction side of theairfoil portion 104. - In
FIG. 5 , the pressure sideperipheral microcircuit 100 is shown. In this circuit, thefirst leg 102 has aninlet 103 which receives cooling fluid from a source (not shown). Theleg 102 provides a flow of cooling fluid which quenches the hot spot in the tip-to-mid region of theairfoil portion 104 shown inFIG. 4B . The cooling fluid within theleg 102 proceeds around a 180degree bend 106 which is supplemented with a plurality offilm holes 108, preferably three film holes. Thefilm holes 108 ensure flow acceleration through thebend 106 to a seconddownstream leg 110 which ends below theplatform 112 of theturbine engine component 90 in anexit 164. Cooling fluid from theleg 110 is fed into an internaltrailing edge circuit 114 to be discussed hereinafter via theexit 164 where it is used to further cool theairfoil portion 104. - Referring now to
FIG. 6 , there is shown a peripheralsuction side microcircuit 120. Thecircuit 120 has afirst leg 122 which communicates with a source (not shown) of cooling fluid. In thefirst leg 122, the cooling flow convects heat away from the suction side. Since thecircuit 120 has no film holes, effective cooling may not be done past the external gage point of theairfoil portion 104 where any film cooling would provide high aerodynamic penalties due to mixing. (PLEASE CHECK THIS TO SEE IF IT MAKES SENSE) Thus, thecircuit 120 is used to feed cooling fluid to aleading edge microcircuit 124 which wraps around the leading edge 126 of theairfoil portion 104. Thecircuit 120 feeds or supplies cooling fluid to the leading edge wrap aroundcircuit 124 through a plurality of wall cross over holes 128. As can be seen fromFIG. 6 , thecircuit 120 has abend 130 and asecond leg 132. Theholes 128 are preferably located in the vicinity of thebend 130 and thesecond leg 132. Thesecond leg 132 may also communicate with the wrap aroundcircuit 124 via apassageway 134. As themicrocircuit 124 wraps around the leading edge,several holes 136 are located in the leading edge and are used to cool the leading edge of theairfoil portion 104. Further, themicrocircuit 124 is provided with a plurality of film holes 138 for creating a film of cooling fluid over the pressure side of the airfoil portion. - Referring now to
FIG. 7 , there is shown the main body internal cooling circuits which include a leading edgeinternal cooling circuit 150 and the trailing edgeinternal cooling circuit 114. The leading edgeinternal cooling circuit 150 communicates with a source (not shown) of cooling fluid, such as engine bleed air, via aninlet 151 and has one or more film cooling holes 152 adjacent thetip 154 of theairfoil portion 104 to provide tip cooling. Thecircuit 150 also has a plurality ofcross-over holes 156 for supplying cooling fluid to theleading edge microcircuit 124. - The trailing edge
internal circuit 114 also communicates with a source (not shown) of cooling fluid, such as engine bleed air, via aninlet 157 and has one or more film cooling holes 158 adjacent thetip 154 to provide tip cooling. Thecircuit 114 also has a plurality ofcross-over holes 160 for communicating with a trailingedge cooling circuit 162 for cooling the trailing edge of theairfoil portion 104. As can be seen fromFIG. 7 , the tailing edgeinternal circuit 114 also receives cooling fluid from the peripheralpressure side microcircuit 100 via theexit 164. - Each of the leading edge
internal circuit 150 and the trailing edgeinternal circuit 114 may be provided with a plurality of film cooling holes 170 and 172 respectively to form cooling films over the pressure and suction sides of theairfoil portion 104. - Using the pressure and suction side cooling circuits of the present invention, the airfoil portion of a turbine engine component may be very effectively convectively cooled. Using the pressure side circuit, the cooling flow is returned to the trailing edge internal circuit for further cooling of the airfoil. Using the suction side circuit, the leading edge of the airfoil is cooled first before discharging in pressure side film. This effective use of coolant allows for positive effects on cycle thermodynamic efficiency, turbine efficiency, rotor inlet temperature impacts, and specific fuel consumption.
- It is apparent that there has been provided in accordance with the present invention serpentine microcircuits for hot gas migration which fully satisfy the objects, means, and advantages set forth hereinbefore. While the present invention has been described in the context of specific embodiments thereof, other unforeseeable alternatives, modifications, and variations may become apparent to those skilled in the art having read the foregoing description. Accordingly, it is intended to embrace those alternatives, modifications, and variations as fall within the broad scope of the appended claims.
Claims (17)
Priority Applications (3)
Application Number | Priority Date | Filing Date | Title |
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US11/494,831 US7581928B1 (en) | 2006-07-28 | 2006-07-28 | Serpentine microcircuits for hot gas migration |
EP20070252841 EP1881157B1 (en) | 2006-07-18 | 2007-07-18 | Serpentine microcircuits for local heat removal |
JP2007194055A JP2008032008A (en) | 2006-07-28 | 2007-07-26 | Serpentine microcircuit for transferring high temperature gas |
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US11/494,831 US7581928B1 (en) | 2006-07-28 | 2006-07-28 | Serpentine microcircuits for hot gas migration |
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US20090208343A1 true US20090208343A1 (en) | 2009-08-20 |
US7581928B1 US7581928B1 (en) | 2009-09-01 |
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US11/494,831 Active 2027-11-23 US7581928B1 (en) | 2006-07-18 | 2006-07-28 | Serpentine microcircuits for hot gas migration |
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Cited By (3)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20090324385A1 (en) * | 2007-02-15 | 2009-12-31 | Siemens Power Generation, Inc. | Airfoil for a gas turbine |
US20130236329A1 (en) * | 2012-03-09 | 2013-09-12 | United Technologies Corporation | Rotor blade with one or more side wall cooling circuits |
US10415394B2 (en) * | 2013-12-16 | 2019-09-17 | United Technologies Corporation | Gas turbine engine blade with ceramic tip and cooling arrangement |
Families Citing this family (3)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US9422817B2 (en) | 2012-05-31 | 2016-08-23 | United Technologies Corporation | Turbine blade root with microcircuit cooling passages |
US11143038B2 (en) | 2013-03-04 | 2021-10-12 | Raytheon Technologies Corporation | Gas turbine engine high lift airfoil cooling in stagnation zone |
US10731469B2 (en) | 2016-05-16 | 2020-08-04 | Raytheon Technologies Corporation | Method and apparatus to enhance laminar flow for gas turbine engine components |
Citations (7)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US3191908A (en) * | 1961-05-02 | 1965-06-29 | Rolls Royce | Blades for fluid flow machines |
US5348446A (en) * | 1993-04-28 | 1994-09-20 | General Electric Company | Bimetallic turbine airfoil |
US5667359A (en) * | 1988-08-24 | 1997-09-16 | United Technologies Corp. | Clearance control for the turbine of a gas turbine engine |
US5813835A (en) * | 1991-08-19 | 1998-09-29 | The United States Of America As Represented By The Secretary Of The Air Force | Air-cooled turbine blade |
US6036441A (en) * | 1998-11-16 | 2000-03-14 | General Electric Company | Series impingement cooled airfoil |
US20050265837A1 (en) * | 2003-03-12 | 2005-12-01 | George Liang | Vortex cooling of turbine blades |
US7416390B2 (en) * | 2005-03-29 | 2008-08-26 | Siemens Power Generation, Inc. | Turbine blade leading edge cooling system |
Family Cites Families (4)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US5931638A (en) | 1997-08-07 | 1999-08-03 | United Technologies Corporation | Turbomachinery airfoil with optimized heat transfer |
US6254334B1 (en) | 1999-10-05 | 2001-07-03 | United Technologies Corporation | Method and apparatus for cooling a wall within a gas turbine engine |
GB0114503D0 (en) | 2001-06-14 | 2001-08-08 | Rolls Royce Plc | Air cooled aerofoil |
US7011502B2 (en) | 2004-04-15 | 2006-03-14 | General Electric Company | Thermal shield turbine airfoil |
-
2006
- 2006-07-28 US US11/494,831 patent/US7581928B1/en active Active
-
2007
- 2007-07-26 JP JP2007194055A patent/JP2008032008A/en active Pending
Patent Citations (7)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US3191908A (en) * | 1961-05-02 | 1965-06-29 | Rolls Royce | Blades for fluid flow machines |
US5667359A (en) * | 1988-08-24 | 1997-09-16 | United Technologies Corp. | Clearance control for the turbine of a gas turbine engine |
US5813835A (en) * | 1991-08-19 | 1998-09-29 | The United States Of America As Represented By The Secretary Of The Air Force | Air-cooled turbine blade |
US5348446A (en) * | 1993-04-28 | 1994-09-20 | General Electric Company | Bimetallic turbine airfoil |
US6036441A (en) * | 1998-11-16 | 2000-03-14 | General Electric Company | Series impingement cooled airfoil |
US20050265837A1 (en) * | 2003-03-12 | 2005-12-01 | George Liang | Vortex cooling of turbine blades |
US7416390B2 (en) * | 2005-03-29 | 2008-08-26 | Siemens Power Generation, Inc. | Turbine blade leading edge cooling system |
Cited By (4)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20090324385A1 (en) * | 2007-02-15 | 2009-12-31 | Siemens Power Generation, Inc. | Airfoil for a gas turbine |
US7871246B2 (en) * | 2007-02-15 | 2011-01-18 | Siemens Energy, Inc. | Airfoil for a gas turbine |
US20130236329A1 (en) * | 2012-03-09 | 2013-09-12 | United Technologies Corporation | Rotor blade with one or more side wall cooling circuits |
US10415394B2 (en) * | 2013-12-16 | 2019-09-17 | United Technologies Corporation | Gas turbine engine blade with ceramic tip and cooling arrangement |
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JP2008032008A (en) | 2008-02-14 |
US7581928B1 (en) | 2009-09-01 |
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