EP2392774B1 - Turbinenschaufel mit umfassendem Vorderkantenkühlkanal - Google Patents

Turbinenschaufel mit umfassendem Vorderkantenkühlkanal Download PDF

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Publication number
EP2392774B1
EP2392774B1 EP10005822.1A EP10005822A EP2392774B1 EP 2392774 B1 EP2392774 B1 EP 2392774B1 EP 10005822 A EP10005822 A EP 10005822A EP 2392774 B1 EP2392774 B1 EP 2392774B1
Authority
EP
European Patent Office
Prior art keywords
airfoil
portions
leading edge
cooling
cores
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.)
Not-in-force
Application number
EP10005822.1A
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English (en)
French (fr)
Other versions
EP2392774A1 (de
Inventor
William Abdel-Messeh
Justin D. Piggush
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Raytheon Technologies Corp
Original Assignee
United Technologies Corp
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Filing date
Publication date
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Priority to EP10005822.1A priority Critical patent/EP2392774B1/de
Publication of EP2392774A1 publication Critical patent/EP2392774A1/de
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Publication of EP2392774B1 publication Critical patent/EP2392774B1/de
Not-in-force legal-status Critical Current
Anticipated expiration legal-status Critical

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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
    • 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
    • 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
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/186—Film cooling

Definitions

  • This disclosure relates to a cooling passage for an airfoil.
  • Turbine blades are utilized in gas turbine engines.
  • a turbine blade typically includes a platform having a root on one side and an airfoil extending from the platform opposite the root. The root is secured to a turbine rotor.
  • Cooling circuits are formed within the airfoil to circulate cooling fluid, such as air.
  • multiple relatively large cooling channels extend radially from the root toward a tip of the airfoil. Air flows through the channels and cools the airfoil, which is relatively hot during operation of the gas turbine engine.
  • Some advanced cooling designs use one or more radial cooling passages that extend from the root toward the tip near a leading edge of the airfoil.
  • the cooling passages are arranged between the cooling channels and an exterior surface of the airfoil.
  • the cooling passages provide extremely high convective cooling.
  • Prior art leading edge cooling arrangements typically include two cooling approaches. First, internal impingement cooling is used, which produces high internal heat transfer rates. Second, showerhead film cooling is used to create a film on the external surface of the airfoil. Relatively large amounts of cooling flow are required, which tends to exit the airfoil at relatively cool temperatures. The heat that the cooling flow absorbs is relatively small since the cooling flow travels along short paths within the airfoil, resulting in cooling inefficiencies.
  • a cooling passage wrapped at the leading edge is a cooling passage wrapped at the leading edge.
  • This wrapped leading edge cooling passage is formed by a refractory metal core that is secured to another core.
  • the cores are placed in a mold, and a superalloy is cast into the mold about the cores to form the airfoil.
  • the cores are removed from the cast airfoil to provide the cooling passages.
  • the wrapped leading edge cooling passage does not provide the amount of desired cooling to the leading edge.
  • a turbine engine airfoil having multiple cooling passages formed in the leading edge of the airfoil is disclosed in EP-A-1467064 .
  • a turbine engine airfoil in accordance with the invention is set forth in claim 1.
  • the second cooling passage extends radially, and the first cooling passage wraps around a portion of the second cooling passage from a pressure side to a suction side between the second cooling passage and the exterior surface.
  • the first portion is arranged between the pressure and suction sides.
  • the invention also provides a method of manufacturing an airfoil, as set forth in claim 8.
  • Figure 1 schematically illustrates a gas turbine engine 10 that includes a fan 14, a compressor section 16, a combustion section 18 and a turbine section 11, which are disposed about a central axis 12.
  • air compressed in the compressor section 16 is mixed with fuel that is burned in combustion section 18 and expanded in the turbine section 11.
  • the turbine section 11 includes, for example, rotors 13 and 15 that, in response to expansion of the burned fuel, rotate, which drives the compressor section 16 and fan 14.
  • the turbine section 11 includes alternating rows of blades 20 and static airfoils or vanes 19. It should be understood that Figure 1 is for illustrative purposes only and is in no way intended as a limitation on this disclosure or its application.
  • FIG. 2 An example blade 20 is shown in Figure 2 .
  • the blade 20 includes a platform 32 supported by a root 36, which is secured to a rotor.
  • An airfoil 34 which falls outside the scope of the invention, extends radially outwardly from the platform 32 opposite the root 36. While the airfoil 34 is disclosed as being part of a turbine blade 20, it should be understood that the disclosed airfoil can also be used as a vane.
  • the airfoil 34 includes an exterior surface 57 extending in a chord-wise direction C from a leading edge 38 to a trailing edge 40.
  • the airfoil 34 extends between pressure and suction sides 42, 44 in a airfoil thickness direction T, which is generally perpendicular to the chord-wise direction C.
  • the airfoil 34 extends from the platform 32 in a radial direction R to an end portion or tip 33.
  • a cooling trench 48 is provided on the leading edge 38 to create a cooling film on the exterior surface 57. In the examples, the trench 48 is arranged in proximity to a stagnation line on the leading edge 38, which is an area in which there is little or no fluid flow over the leading edge.
  • FIG 3A schematically illustrates an airfoil molding process in which a mold 94 having mold halves 94A, 94B provides a mold contour that defines the exterior surface 57 of the airfoil 34.
  • cores 82 which may be ceramic, are arranged within the mold 94 to provide the cooling channels 50, 52, 54 ( Figure 3C ).
  • cores 82 which may be ceramic, are arranged within the mold 94 to provide the cooling channels 50, 52, 54 ( Figure 3C ).
  • multiple, relatively large radial cooling channels 50, 52, 54 are provided internally within the airfoil 34 to deliver airflow for cooling the airfoil.
  • the cooling channels 50, 52, 54 typically provide cooling air from the root 36 of the blade 20.
  • the airfoil 34 includes a first cooling passage 56 arranged near the leading edge 38.
  • the first cooling passage 56 is in fluid communication with the cooling channel 50, in the example shown.
  • One or more core structures 68 ( Figures 3A and 3B ), such as refractory metal cores, are arranged within the mold 94 and connected to the other cores 82.
  • the core structure 68 which is generally C-shaped, provides the first cooling passage 56 in the example disclosed.
  • the core structure 68 (shown in Figure 3B ) is stamped from a flat sheet of refractory metal material. The core structure 68 is then bent or shaped to a desired contour.
  • the ceramic core and/or refractory metal cores are removed from the airfoil 34 after the casting process by chemical or other means.
  • a core assembly can be provided in which a portion of the core structure 68 is received in a recess of the other core 82, as shown in Figure 3A .
  • the resultant first cooling passage 56 provided by the core structure 68 is in fluid communication with the cooling channel 50 subsequent to the airfoil casting process.
  • the core structure 68 includes a first portion 72 and a second portion.
  • the second portion includes multiple, radially spaced first and second sets of arcuate legs 74, 76 that wrap around a portion of the cooling channel 50.
  • the shape of the legs 74, 76 generally mirror the exterior surface 57 of the leading edge 38.
  • the first and second sets of legs 74, 76 are secured to the other core 82.
  • One set of legs 74 is arranged on the pressure side 42 and the other set of legs 76 is arranged on the suction side 44.
  • the first portion 72 does not extend to the exterior surface 57.
  • the trench 48 is formed by a chemical or mechanical machining process, for example, to fluidly connect the first portion 72 to the leading edge 38. Cooling fluid is provided from the first cooling channel 50 through the first cooling passage 56 to provide a cooling film on the leading edge 38 via the trench 48.
  • a core structure 168 for a further airfoil which falls outside the scope of the invention is shown that provides the trench 48 during the casting process.
  • the first portion 172 extends beyond the exterior surface and into the mold 94 where the first portion 172 is held by a core retention feature 96, which is provided by a notch in the mold 94, for example.
  • a trench will be provided at the leading edge 138.
  • the legs 174, 176 are at an angle or transverse laterally to the first portion 172.
  • the example core structure 168 provides first and second sets of legs 174, 176 on opposite sides and in radially spaced, alternating relationship from one another.
  • the first portion 172 extends in a direction opposite the other core 82.
  • the first cooling passage is provided by multiple separate networks of passageways, as illustrated in Figures 5A and 5B .
  • the networks of passageways are formed with multiple core structures 86, 88 having first portions 272, 273 that are discrete from one another.
  • One of the cores structures 86 is arranged on the suction side 44 and the other core structure 88 is arranged on the pressure side 42.
  • the legs 274, 276 are only fluidly connected to one another through the cooling channel 50.
  • the first portions 272, 273 extend beyond the exterior surface 57 in the leading edge 238 and are configured to provide laterally spaced trenches 248 on the airfoil 234, as shown in Figure 5C .
  • the trenches 248 may be laterally and radially staggered.
  • the first cooling passage is provided by two networks of passageways created by core structures 186a, 186b, 188a, 188b provided on each of the pressure and suction sides 42, 44 of airfoil 334.
  • the core structures 186a, 186b, 188a, 188b respectively provide discrete first portions 273a, 273b, 272a, 272b that create trenches 348 in leading edge 338, shown in Figure 6B .

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Molds, Cores, And Manufacturing Methods Thereof (AREA)

Claims (12)

  1. Turbinenschaufel (34), umfassend:
    eine Schaufelstruktur, die eine Außenfläche (57) beinhaltet, die eine Vorderkante (38) bereitstellt, mehrere sich radial erstreckende erste Kühlkanäle in der Nähe der Vorderkante (38), von denen jeder einen ersten Abschnitt und zweite Abschnitte beinhaltet, wobei sich der erste Abschnitt zur Außenfläche (57) hin erstreckt und einen sich radial erstreckenden Graben (248; 348) in der Vorderkante (38) bildet und die zweiten Abschnitte in Fluidkommunikation mit einem zweiten Kühlkanal (50) stehen und radial beabstandete gebogene Beine sind, wobei die sich radial erstreckenden Gräben (248; 348) der mehreren ersten Kühlkanäle seitlich beabstandet sind.
  2. Turbinenschaufel nach Anspruch 1, wobei der zweite Kühlkanal (50) sich radial erstreckt und der erste Kühlkanal um einen Abschnitt des zweiten Kühlkanals (50) von einer Druckseite bis zu einer Ansaugseite zwischen dem zweiten Kühlkanal und der Außenfläche (57) herum verläuft, wobei der erste Abschnitt zwischen der Druck- und der Ansaugseite angeordnet ist.
  3. Turbinenschaufel nach Anspruch 2, wobei der erste Kühlkanal im Wesentlichen C-förmig ist.
  4. Turbinenschaufel nach Anspruch 1, wobei sich einer der mehreren ersten Kühlkanäle auf der Druckseite befindet und sich ein anderer der Kanäle auf der Ansaugseite befindet, wobei jeder der zweiten Abschnitte des ersten Kühlkanals ausschließlich durch den zweiten Kühlkanal (50) mit den zweiten Abschnitten anderer der mehreren ersten Kühlkanäle fluidverbunden ist.
  5. Turbinenschaufel nach Anspruch 4, wobei mindestens zwei erste Kanäle auf mindestens einer der Druck- und der Ansaugseite angeordnet sind.
  6. Turbinenschaufel nach Anspruch 5, wobei die radial beabstandeten gebogenen Beine in einer alternierenden Beziehung zueinander angeordnet sind.
  7. Turbinenschaufel nach einem der vorhergehenden Ansprüche, wobei die seitlich beabstandeten Gräben (248) zudem radial beabstandet sind.
  8. Verfahren zum Herstellen einer Schaufel mit internen Kühlkanälen, wobei das Verfahren die folgenden Schritte umfasst:
    Bereitstellen mehrerer erster Kerne (186a, 186b, 188a, 188b), von denen jeder einen ersten, sich radial erstreckenden Abschnitt (272, 273; 272a, 272b, 273a, 273b) und mehrere im Wesentlichen gebogene zweite Abschnitte (274, 276), die sich im Wesentlichen in Sehnenrichtung aus dem ersten Abschnitt (272, 273; 272a, 272b, 273a, 273b) heraus erstrecken, aufweist, wobei die zweiten Abschnitte radial zueinander beabstandete bogenförmige Beine sind;
    Anordnen der ersten Kerne (186a, 186b, 188a, 188b) in einer Gussform (94) an einer Stelle, die einer Vorderkante (38) einer Schaufel, die durch die Gussform (94) gebildet werden soll, entspricht, wobei die Gussform (94) eine derartige Schaufelkontur bereitstellt, dass die ersten sich radial erstreckenden Abschnitte (272, 273; 272a, 272b, 273a, 273b) seitlich voneinander beabstandet sind;
    Anordnen eines zweiten Kerns (82) radial in der Gussform (94), wobei der zweite Kern (82) die zweiten Abschnitte (274, 276) der ersten Kerne (186a, 186b, 188a, 188b) stützt; und
    Aufbringen von Gussmaterial in die Gussform (94), wobei die ersten Abschnitte (272, 273; 272a, 272b, 273a, 273b), die sich über die Schaufelkontur hinaus in die Gussform (94) hinein erstrecken, und die zweiten Abschnitte (274, 276) von dem Gussmaterial umgeben sind, wobei die ersten Abschnitte (272, 273; 272a, 272b, 273a, 273b) mehreren seitlich beabstandeten und sich radial erstreckenden Gräben (48) in der Vorderkante (38) entsprechen.
  9. Verfahren nach Anspruch 8, umfassend den Schritt des Haltens der ersten Abschnitte (272, 273; 272a, 272b, 273a, 273b) in der Gussform (94) in einer Kernrückhaltefunktion, wobei sich der erste Abschnitt (272, 273; 272a, 272b, 273a, 273b) außerhalb des Gussmaterials befindet.
  10. Verfahren nach Anspruch 8 oder 9, wobei die ersten Kerne (186a, 186b, 188a, 188b) mindestens ein Kernstück beinhalten, wobei das mindestens eine Kernstück (168) um die Vorderkante (38) verläuft und dabei die Schaufelkontur zwischen Seiten, die der Druck- und der Ansaugseite der Schaufel entsprechen, spiegelt.
  11. Verfahren nach einem der Ansprüche 8 bis 10, wobei der zweite Kern (82) ein Keramikkern ist und die ersten Kerne (186a, 186b, 188a, 188b) hochschmelzende Metallkerne sind, wobei die ersten und die zweiten Kernstücke (186a, 186b, 188a, 188b, 82) aneinander befestigt sind.
  12. Verfahren nach einem der Ansprüche 8 bis 11, wobei die ersten Kerne (186a, 186b, 188a, 188b) durch Stanzen einer Kernstruktur, einschließlich einer gewünschten Form, aus einem hochschmelzenden metallischen Material und Biegen der ersten Kernstücke (86, 88; 186a, 186b, 188a, 188b), um eine gewünschte Kontur bereitzustellen, bereitgestellt werden.
EP10005822.1A 2010-06-04 2010-06-04 Turbinenschaufel mit umfassendem Vorderkantenkühlkanal Not-in-force EP2392774B1 (de)

Priority Applications (1)

Application Number Priority Date Filing Date Title
EP10005822.1A EP2392774B1 (de) 2010-06-04 2010-06-04 Turbinenschaufel mit umfassendem Vorderkantenkühlkanal

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Application Number Priority Date Filing Date Title
EP10005822.1A EP2392774B1 (de) 2010-06-04 2010-06-04 Turbinenschaufel mit umfassendem Vorderkantenkühlkanal

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EP2392774A1 EP2392774A1 (de) 2011-12-07
EP2392774B1 true EP2392774B1 (de) 2019-03-06

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Families Citing this family (13)

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Publication number Priority date Publication date Assignee Title
WO2014126674A1 (en) * 2013-02-12 2014-08-21 United Technologies Corporation Gas turbine engine component cooling passage and space eating core
US10240464B2 (en) 2013-11-25 2019-03-26 United Technologies Corporation Gas turbine engine airfoil with leading edge trench and impingement cooling
US9987677B2 (en) 2015-12-17 2018-06-05 General Electric Company Method and assembly for forming components having internal passages using a jacketed core
US10099276B2 (en) 2015-12-17 2018-10-16 General Electric Company Method and assembly for forming components having an internal passage defined therein
US10150158B2 (en) 2015-12-17 2018-12-11 General Electric Company Method and assembly for forming components having internal passages using a jacketed core
US9579714B1 (en) 2015-12-17 2017-02-28 General Electric Company Method and assembly for forming components having internal passages using a lattice structure
US10137499B2 (en) 2015-12-17 2018-11-27 General Electric Company Method and assembly for forming components having an internal passage defined therein
US10046389B2 (en) 2015-12-17 2018-08-14 General Electric Company Method and assembly for forming components having internal passages using a jacketed core
US10118217B2 (en) 2015-12-17 2018-11-06 General Electric Company Method and assembly for forming components having internal passages using a jacketed core
US10099283B2 (en) 2015-12-17 2018-10-16 General Electric Company Method and assembly for forming components having an internal passage defined therein
US10099284B2 (en) 2015-12-17 2018-10-16 General Electric Company Method and assembly for forming components having a catalyzed internal passage defined therein
US10335853B2 (en) 2016-04-27 2019-07-02 General Electric Company Method and assembly for forming components using a jacketed core
US10286450B2 (en) 2016-04-27 2019-05-14 General Electric Company Method and assembly for forming components using a jacketed core

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US5486093A (en) * 1993-09-08 1996-01-23 United Technologies Corporation Leading edge cooling of turbine airfoils
US6164912A (en) * 1998-12-21 2000-12-26 United Technologies Corporation Hollow airfoil for a gas turbine engine
US6955522B2 (en) * 2003-04-07 2005-10-18 United Technologies Corporation Method and apparatus for cooling an airfoil
US7306026B2 (en) * 2005-09-01 2007-12-11 United Technologies Corporation Cooled turbine airfoils and methods of manufacture
US8572844B2 (en) * 2008-08-29 2013-11-05 United Technologies Corporation Airfoil with leading edge cooling passage

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