WO2018080416A1 - Profil aérodynamique de turbine avec passages de paroi proche sans nervures de liaison - Google Patents

Profil aérodynamique de turbine avec passages de paroi proche sans nervures de liaison Download PDF

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Publication number
WO2018080416A1
WO2018080416A1 PCT/US2016/058365 US2016058365W WO2018080416A1 WO 2018080416 A1 WO2018080416 A1 WO 2018080416A1 US 2016058365 W US2016058365 W US 2016058365W WO 2018080416 A1 WO2018080416 A1 WO 2018080416A1
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WO
WIPO (PCT)
Prior art keywords
radial
coolant channel
hollow rib
airfoil
turbine airfoil
Prior art date
Application number
PCT/US2016/058365
Other languages
English (en)
Inventor
Paul A. SANDERS
Denis TORRES GONZALEZ
Matthew J. GOLSEN
Original Assignee
Siemens Aktiengesellschaft
Priority date (The priority date 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 date listed.)
Filing date
Publication date
Application filed by Siemens Aktiengesellschaft filed Critical Siemens Aktiengesellschaft
Priority to PCT/US2016/058365 priority Critical patent/WO2018080416A1/fr
Publication of WO2018080416A1 publication Critical patent/WO2018080416A1/fr

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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
    • 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
    • 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
    • F01D5/188—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall
    • 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
    • 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
    • F05D2230/22—Manufacture essentially without removing material by sintering
    • 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
    • F05D2230/23—Manufacture essentially without removing material by permanently joining parts together
    • F05D2230/232—Manufacture essentially without removing material by permanently joining parts together by welding
    • 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
    • F05D2230/23—Manufacture essentially without removing material by permanently joining parts together
    • F05D2230/232—Manufacture essentially without removing material by permanently joining parts together by welding
    • F05D2230/237—Brazing

Definitions

  • the present invention is directed generally to turbine airfoils, and more particularly to turbine airfoils having internal cooling channels for conducting a cooling fluid through the airfoil.
  • a turbomachine such as a gas turbine engine
  • air is pressurized in a compressor section and then mixed with fuel and burned in a combustor section to generate hot combustion gases.
  • the hot combustion gases are expanded within a turbine section of the engine where energy is extracted to power the compressor section and to produce useful work, such as turning a generator to produce electricity.
  • the hot combustion gases travel through a series of turbine stages within the turbine section.
  • a turbine stage may include a row of stationary airfoils, i.e., vanes, followed by a row of rotating airfoils, i.e., turbine blades, where the turbine blades extract energy from the hot combustion gases for providing output power. Since the airfoils, i.e., vanes and turbine blades, are directly exposed to the hot combustion gases, they are typically provided with internal cooling channels that conduct a cooling fluid, such as compressor bleed air, through the airfoil.
  • a cooling fluid such as compressor bleed air
  • One type of airfoil extends from a radially inner platform at a root end to a radially outer portion of the airfoil, and includes opposite pressure and suction sidewalk extending span-wise along a radial direction and extending axially from a leading edge to a trailing edge of the airfoil.
  • the cooling channels extend inside the airfoil between the pressure and suction sidewalls and may conduct the cooling fluid in alternating radial directions through the airfoil. The cooling channels remove heat from the pressure sidewall and the suction sidewall and thereby avoid overheating of these parts.
  • aspects of the present invention provide a turbine airfoil with near wall passages without connecting ribs.
  • a turbine airfoil comprising a generally hollow airfoil body formed by an outer wall extending span-wise along a radial direction.
  • the outer wall comprises a pressure sidewall and a suction sidewall joined at a leading edge and a trailing edge.
  • a chordal axis is defined extending generally centrally between the pressure sidewall and the suction sidewall.
  • a plurality of radially extending partition walls positioned in an interior of the airfoil body, connecting the pressure sidewall and the suction sidewall. The partition walls are spaced along the chordal axis to define radial cavities therebetween. At least one radially extending hollow rib is positioned in one of the radial cavities.
  • the hollow rib is spaced from the pressure sidewall and from the suction sidewall to define a first and a second near wall passage respectively.
  • the hollow rib is further spaced from a respective adjacent partition wall on chord-wise opposite sides to define a first and a second connecting passage respectively.
  • the hollow rib forms a first radial coolant channel therewithin, which is in communication with a coolant source external to the turbine airfoil.
  • the hollow rib is mechanically connected to the turbine airfoil at a base of the hollow rib, without any mechanical connection with the pressure and suction sidewalk.
  • first and second near wall passages and the first and second connecting passages are interconnected along a radial extent to form a second radial coolant channel, which surrounds the first radial coolant channel and is fluidically connected in series with the first radial coolant channel.
  • FIG. 1 is a perspective view of a turbine airfoil featuring embodiments of the present invention
  • FIG. 2 is a cross-sectional view through the turbine airfoil along the section ⁇ - ⁇ of FIG. 1 according a first embodiment of the invention
  • FIG 3. is a schematic cross-sectional view along the section ⁇ - ⁇ in FIG. 2 illustrating an exemplary configuration featuring an uncapped hollow rib;
  • FIG. 4 is a schematic cross-sectional view along the section IV-IV in FIG. 2 illustrating an alternate exemplary configuration featuring a hollow rib with an end cap provided with impingement orifices;
  • FIG. 5 illustrates a perspective view of a portion of a casting core for manufacturing an inventive turbine airfoil according to one embodiment of the present invention
  • FIG. 6 illustrates, in perspective view, a radially sliced section along the sectional plane VI- VI, of the portion of the casting core shown in FIG. 5;
  • FIG. 7 illustrates, in perspective view, a circumferentially sliced section along the sectional plane VII- VII, of the portion of the casting core shown in FIG. 5;
  • FIG. 8 is a cross-sectional view through the turbine airfoil along the section VIII-VIII of FIG. 1 according a second embodiment of the invention featuring a hollow rib with impingement orifices along a radial extent;
  • FIG. 9 is a schematic cross-sectional view along the section ⁇ - ⁇ in FIG. 8 illustrating an exemplary configuration featuring a hollow rib with an end cap;
  • FIG 10. is a schematic cross-sectional view along the section X-X in FIG. 8 illustrating an alternate exemplary configuration featuring a hollow rib with an end cap provided with impingement orifices.
  • the direction X denotes an axial direction parallel to an axis of the turbine engine
  • the directions R and T respectively denote a radial direction and a tangential (or circumferential) direction with respect to said axis of the turbine engine.
  • aspects of the present invention relate to an internally cooled turbine airfoil.
  • coolant supplied to the internal cooling channels in a turbine airfoil often comprises air diverted from a compressor section. Achieving a high cooling efficiency based on the rate of heat transfer is a significant design consideration in order to minimize the flow rate of coolant air diverted from the compressor for cooling.
  • Many turbine blades and vanes involve a two-wall structure including a pressure sidewall and a suction sidewall joined at a leading edge and at a trailing edge.
  • Internal cooling channels are created by employing internal partition walls or ribs which connect the pressure and suction sidewalls in a direct linear fashion.
  • near wall cooling To address the problem of efficiently utilizing coolant for targeted convective heat transfer with the airfoil outer wall, techniques have been developed to implement near wall cooling, such as that disclosed in the International Application No. PCT US2015/047332, filed by the present applicant, and herein incorporated by reference in its entirety. Briefly, such a near wall cooling technique employs the use of a flow displacement element in the form of an elongated hollow body that is connected to the airfoil pressure and suction sidewalk along a radial extent by connector ribs. The elongated hollow body defines an inactive cavity therein, through which there is no coolant flow.
  • the flow cross-sectional area of the coolant flow is thus reduced by displacing much of the coolant toward the hot pressure and suction sidewalls, while also increasing the target wall velocities as a result of the narrowing of the flow cross-section.
  • the above results in an increase in convective heat transfer between the coolant and the hot pressure and suction sidewalls. Furthermore, this leads to an efficient use of the coolant as the coolant flow is displaced from the center of the flow cross-section toward the hot walls that need the most cooling, namely, the pressure and suction sidewalls.
  • the present inventors have devised an improvement to the above described configuration which achieves said cooling benefits while further reducing thermal fight between the hot pressure and suction sidewalls and the relatively cold hollow elongated body. This is achieved by supporting the hollow elongated body at the base of the hollow elongated body, while eliminating the connector ribs of the above described configuration.
  • the airfoil 10 is illustrated according to one embodiment.
  • the airfoil 10 is a turbine blade for a gas turbine engine. It should however be noted that aspects of the invention could additionally be incorporated into stationary vanes in a gas turbine engine.
  • the airfoil 10 may include an outer wall 14 adapted for use, for example, in a high pressure stage of an axial flow gas turbine engine.
  • the outer wall 14 extends span-wise along a radial direction R of the turbine engine and includes a generally concave shaped pressure sidewall 16 and a generally convex shaped suction sidewall 18.
  • the pressure sidewall 16 and the suction sidewall 18 are joined at a leading edge 20 and at a trailing edge 22.
  • the outer wall 14 may be coupled to a root 56 at a platform 58.
  • the root 56 may couple the turbine airfoil 10 to a disc (not shown) of the turbine engine.
  • the outer wall 14 is delimited in the radial direction by a radially outer end face or airfoil tip 52 and a radially inner end face 54 coupled to the platform 58.
  • the airfoil 10 may be a stationary turbine vane with a radially inner end face coupled to the inner diameter of the turbine section of the turbine engine and a radially outer end face coupled to the outer diameter of the turbine section of the turbine engine.
  • the outer wall 14 delimits an airfoil interior 11 comprising internal cooling channels, which may receive a coolant, such as air diverted from a compressor section (not shown), via one or more coolant supply passages (not shown) through the root 56.
  • a plurality of partition walls 24 are positioned spaced apart in the along a chordal axis 30. As shown in FIG. 2, the chordal axis 30 is defined extending centrally between the pressure and suction sidewalls 16, 18.
  • the partition walls 24 extend along a radial extent, connecting the pressure sidewall 16 and the suction sidewall 18 to define internal radial cavities 40.
  • At least some of the radial cavities 40 define coolant channels therein, which are individually identified as A, B, C, D, E, F.
  • the coolant traverses through the radial coolant channels A-F, absorbing heat from the airfoil components, particularly the hot outer wall 14.
  • the coolant channels A-F lead the coolant to a leading edge coolant cavity LEC adjacent to the leading edge 20 and to a trailing edge coolant cavity TEC adjacent to the trailing edge 22. From the cavities LEC and TEC, the coolant exits the airfoil 10 via exhaust orifices 27 and 29 positioned along the leading edge 20 and the trailing edge 22 respectively.
  • the exhaust orifices 27 may provide film cooling along the leading edge 20 (see FIG 1). Although not shown in the drawings, film cooling orifices may be provided at multiple locations, including anywhere on the pressure sidewall 16, suction sidewall 18, leading edge 20 and the airfoil tip 52.
  • the turbine airfoil 10 comprises at least one radially extending hollow rib 26 positioned in the airfoil interior 11.
  • two such hollow ribs 26 are provided.
  • Each hollow rib 26 is positioned in one of the radial cavities 40.
  • each hollow rib 26 is spaced from the pressure sidewall 16 and from the suction sidewall 18 to define a first near wall passage 72 and a second near wall passage 74 respectively.
  • Each hollow rib 26 is further spaced from a respective adjacent partition wall 24 on chord- wise opposite sides to define a first connecting passage 76 and a second connecting passage 78 respectively.
  • the near wall passages 72, 74 and the connecting passages 76, 78 extend along the radial direction, with the connecting passages 76, 78 extending transversely between the near wall passages 72, 74.
  • each hollow rib 26 comprises first and second opposite side faces 42 and 44.
  • the first side face 42 is spaced from the pressure sidewall 16 such that the first near wall passage 72 is defined between the first side face 42 and the pressure sidewall 16.
  • the second side face 44 is spaced from the suction sidewall 18 such that the second near wall passage 74 is defined between the second side face 44 and the suction sidewall 18.
  • Each hollow rib 26 further comprises third and fourth opposite side faces 46 and 48 extending between the first and second side faces 42 and 44.
  • the third and fourth side faces 46 and 48 are respectively spaced from the partition walls 24 on either side to define a respective connecting passage 76, 78 between the respective side face 46, 48 and the respective partition wall 24.
  • FIG. 3 shows a cross-sectional view of a first of the hollow ribs 26, namely the forward located hollow rib 26.
  • the hollow rib 26 forms a first radial coolant channel B, which is in communication with a coolant source external to the turbine airfoil 10, such as air diverted from a compressor section (not shown).
  • the hollow rib 26 is mechanically connected to the turbine airfoil 10 at a base 50 of the hollow rib 26.
  • the mechanical connection may be located at the level of or underneath (radially inboard of) the inner diameter endwall/platform 58 of the turbine airfoil 10, which in this example is a turbine blade.
  • first and second near wall passages 72, 74 and the first and second connecting passages 76, 78 are interconnected along a radial extent to form a respective second radial coolant channel C.
  • the second radial coolant channel C surrounds the first radial coolant channel B and is fluidically connected in series with the first radial coolant channel B (see FIG. 3).
  • the second or aft located hollow rib 26 which defines a first radial coolant channel D therewithal.
  • the second hollow rib 26 may also be connected to the turbine airfoil 10 at a base of the second hollow rib 26, without any mechanical connection with the pressure and suction sidewalls 16, 18, such that the first and second near wall passages 72, 74 and the first and second connecting passages 76, 78 are interconnected along a radial extent to form a respective second radial coolant channel E, which surrounds the respective first radial coolant channel D and is fluidically connected in series with the respective first radial coolant channel D.
  • the hollow ribs 26 may be cast integrally with the airfoil 10, for example from a ceramic casting core.
  • Other manufacturing techniques may include, for example, additive manufacturing processes such as 3-D printing. This allows the inventive aspects to be used for highly contoured airfoils, including 3-D contoured blades and vanes.
  • other manufacturing techniques are within the scope of the present invention, including, for example, assembly (via welding, brazing, etc.) or plastic forming, among others.
  • FIG. 5-7 illustrate a portion of a casting core 150 for manufacturing an inventive turbine airfoil according to one embodiment of the present invention.
  • the portion of the casting core 150 depicted herein corresponds to the portion 200 of the manufactured airfoil shown in FIG. 2.
  • FIG. 5 illustrates a perspective view of the core 150
  • FIG. 6 and 7 respectively illustrate radial and circumferential sliced sections.
  • the core 150 is formed of a plurality of core elements, wherein: the core element 152 forms the first radial coolant channel D (see FIG. 2); the core element 154 forms the surrounding second radial coolant channel E (see FIG. 2); the gap 156 between the core elements 152 and 154 forms the radially extending hollow rib 26 (see FIG.
  • the core element 158 forms the chord- wise adjacent radial coolant channel F (see FIG. 2); the connecting core element 164, which connects the core elements 154 and 158, forms a flow turning passage that fluidically connects the radial coolant channel E with the radial coolant channel F (see FIG. 2); and the core element 160 forms the trailing edge cavity TEC (see FIG. 2).
  • the legs 172, 174 represent core feeds.
  • the present design provides that the coolant air feeds into the hollow rib 26 directly from the coolant source, whereby fewer core feeds are required to enter the aii foil than in the case where the hollow rib 26 would be used as an inactive cavity Mechanical connection at the platform or root of the turbine blade is thereby strengthened.
  • the outer wall 14 including the pressure and suction sidewalk 16, 18 is exposed to hot combustion gases and is therefore at a much higher temperature than internal components, such as the hollow ribs 26.
  • Elimination of a mechanical connection between the hollow ribs 26 and the pressure and suction sidewalk 16, 18 significantly reduces thermal fight between the hollow ribs 26 and the pressure and suction sidewalk 16, 18 along the radial direction.
  • the first radial coolant channels B, D defined within the respective hollow ribs 26, serve as active cavities which conduct coolant flow.
  • the first radial coolant channel B and the second radial coolant channel C conduct a coolant K in opposite radial directions to form an aft-to -forward serpentine cooling path.
  • the first radial coolant channel B is configured as an "up" pass conducting coolant K in a radially outboard direction
  • the second radial coolant channel C is configured as a "down" pass conducting coolant K in a radially radially inboard direction.
  • the hollow rib 26 comprises a first radial end 62 proximate to which the base 50 of the hollow rib 26 is located, and a second radial end 64 opposite to the first radial end 62.
  • the first radial coolant channel B comprises an inlet 82 at the first radial end 62 and an outlet 84 at the second radial end 64.
  • the first radial coolant cavity B is uncapped or substantially open at the second radial end 64.
  • the inlet 82 receives the coolant K from a coolant source external to the turbine airfoil, which may be air diverted from a compressor section (not shown).
  • the coolant K flows in a radially outboard direction to the outlet 84, which is fluidically connected in series to the surrounding narrow-width second radial coolant channel C via a flow turning passage 86.
  • the flow turning passage 86 may be located between the second radial end 64 of the hollow rib 26 and a radial end face 52 of the airfoil body 12, which in this case is the airfoil tip.
  • the second radial coolant channel C may be fluidically connected in series to a chord-wise adjacent radial coolant channel A via a flow turning passage 88.
  • the radial coolant channel A may form part of a leading edge cooling circuit in communication with exhaust orifices 27 located at the leading edge 20 of the airfoil 10 (see FIG. 2).
  • the coolant K may enter the leading edge coolant cavity LEC, for example, via impingement openings provided on the intervening partition wall 24, and then be discharged into the hot gas path via exhaust orifices 27 on the outer wall which may collectively form a shower head for cooling the leading edge 20 of the airfoil 10.
  • the first radial coolant channel D and the second radial coolant channel E may conduct a coolant in opposite radial directions to form a forward-to-aft serpentine cooling path.
  • the first radial coolant channel D is configured as an "up" pass conducting coolant in a radially outboard direction
  • the second radial coolant channel E is configured as a "down" pass conducting coolant in a radially inboard direction.
  • the first radial coolant channel D may comprise an inlet at a first radial end of the hollow rib 26 located near the base of the hollow rib 26.
  • An outlet of the first radial coolant channel D may be located at a second radial end of the hollow rib 26 opposite to the first radial end.
  • the inlet may be located inboard of the platform 58, while the outlet may be located proximate to the airfoil tip 52.
  • the inlet receives a coolant from a coolant source external to the turbine airfoil, which in this case is air diverted from a compressor section.
  • the coolant flows in a radially outboard direction to the outlet of the first radial coolant channel D, which is fluidically connected in series to the surrounding narrow-width second radial coolant channel E via a flow turning passage located between the second radial end of the hollow rib 26 and a radial end face, which in this case is the airfoil tip 52.
  • the second radial coolant channel E may be fluidically connected to a chord-wise adjacent radial coolant channel F via flow turning passage located radially inboard end thereof.
  • the radial coolant channel E may form part of a trailing edge cooling circuit in communication with exhaust orifices 29 located at the trailing edge 22 of the airfoil 10.
  • the radial coolant channel F may be in fluid communication with the trailing edge coolant cavity TEC, which may incorporate trailing edge cooling features as known to one skilled in the art, for example, comprising turbulators, or pin fins, or combinations thereof, before being discharged into the hot gas path via exhaust orifices 29 located along the trailing edge 22.
  • trailing edge cooling features as known to one skilled in the art, for example, comprising turbulators, or pin fins, or combinations thereof, before being discharged into the hot gas path via exhaust orifices 29 located along the trailing edge 22.
  • FIG. 4 illustrates an alternate configuration to FIG. 3.
  • one or both of the hollow ribs 26 may be provided with a perforated end cap at the downstream end.
  • FIG. 4 illustrates only the forward located rib 26.
  • the first radial coolant channel B is capped via an end cap 90 provided at the second end 64 of the hollow rib 26.
  • the end cap 90 is provided with a plurality of impingement orifices 100 through which the coolant K from the first radial coolant channel B impinges on the airfoil tip 52. Additional cooling of the airfoil tip 52 may thus be achieved with the impingement at the exit of the hollow rib 26.
  • the coolant K turns radially inboard into the surrounding narrow-width radial coolant channel C via the flow turning passage 86 located between the airfoil tip 52 and the second end 64 (radially outboard end) of the hollow rib 26.
  • the remaining aspects of the serpentine cooling circuit may be substantially similar to that described in connection with FIG. 2-3. Additionally or alternately, the exemplary configuration of FIG. 4 may be implemented for the aft located hollow rib 26, by capping the first radial coolant channel D via a perforated end cap at the radially outboard end of the hollow rib 26.
  • FIG. 8-10 illustrate still other embodiments of the present invention, in which the hollow rib 26 may be used to provide impingement cooling to the airfoil.
  • the hollow rib 26 may comprise a plurality of impingement orifices 100A positioned along a radial extent of the hollow rib 26 (see FIG. 10-11).
  • the impingement orifices 100A may be configured such that coolant K from the first radial coolant channel B, D is discharged into the respective second radial coolant channel C, E via impingement on to the pressure sidewall 16 and/or the suction sidewall 18. To this end, as shown in FIG.
  • impingement orifices 100 A may be provided on the first side face 42 and/or the second sides face 44 of the hollow rib 26. Additional impingement orifices (not shown) may also be provided on the third and/or fourth side faces 46, 48 of the hollow rib 26.
  • the remainder of the cooling circuit may be substantially similar to that described above in connection with FIG. 2-4, and will hence not be described further.
  • an end cap 90 may be provided at the second radial end 64 of the hollow rib 26.
  • the first radial coolant channel B is completely capped via the end cap 90.
  • the end cap 90 may be perforated, comprising a plurality of further impingement orifices 100B through which coolant K from the first radial coolant channel B impinges on the airfoil tip 52, to provide additional impingement cooling at the airfoil tip 52.
  • FIG. 2-4 and FIG. 8- 10 may be applied singly or in combination in a given turbine airfoil.
  • the configuration of one or more of the hollow ribs may be altered, for example, by mechanically connecting the hollow rib to the airfoil structure at the level of, or radially outboard from, an outer diameter shroud.
  • the hollow rib would extend radially inward from the base, and the resulting serpentine cooling circuits may be radially inverted in relation to the above illustrated embodiments.

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  • General Engineering & Computer Science (AREA)
  • Architecture (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

L'invention concerne un profil aérodynamique de turbine (10) comprenant une nervure creuse s'étendant radialement (26) et positionnée dans une cavité radiale (40) définie entre des parois de séparation espacées dans le sens de la corde (24). La nervure creuse (26) est espacée d'une paroi latérale de pression (16) et d'une paroi latérale d'aspiration (18) pour définir respectivement des premier et second passages de paroi proche (72, 74), et est en outre espacée d'une paroi de séparation adjacente respective (24) sur des côtés opposés dans le sens de la corde pour définir respectivement des premier et second passages de liaison (76, 78). La nervure creuse (26) forme un premier canal de refroidissement radial (B, D) à l'intérieur, en communication avec une source de fluide de refroidissement. La nervure creuse (26) est reliée mécaniquement au profil aérodynamique de turbine (10) au niveau d'une base (50) de la nervure creuse (26). Les premier et second passages de paroi proche (72, 74) et les premier et second passages de liaison (76, 78) sont interconnectés le long d'une étendue radiale pour former un second canal de liquide de refroidissement radial (C, E) qui entoure le premier canal de liquide refroidissement radial (B, D) et qui est en communication fluidique en série avec ce dernier.
PCT/US2016/058365 2016-10-24 2016-10-24 Profil aérodynamique de turbine avec passages de paroi proche sans nervures de liaison WO2018080416A1 (fr)

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PCT/US2016/058365 WO2018080416A1 (fr) 2016-10-24 2016-10-24 Profil aérodynamique de turbine avec passages de paroi proche sans nervures de liaison

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PCT/US2016/058365 WO2018080416A1 (fr) 2016-10-24 2016-10-24 Profil aérodynamique de turbine avec passages de paroi proche sans nervures de liaison

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Cited By (1)

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EP3597859B1 (fr) * 2018-07-13 2023-08-30 Honeywell International Inc. Aube de turbine avec système de refroidissement tolérant à la poussière

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EP1136652A1 (fr) * 2000-03-23 2001-09-26 General Electric Company Segment d'ailette de guidage à turbine avec circulation interne de refroidissement
US7695245B1 (en) * 2007-03-06 2010-04-13 Florida Turbine Technologies, Inc. Turbine airfoil with a multi-impingement cooled spar and shell
EP2628901A1 (fr) * 2012-02-15 2013-08-21 Siemens Aktiengesellschaft Aube de turbine à gaz avec refroidissement par impact
US20140075947A1 (en) * 2012-09-18 2014-03-20 United Technologies Corporation Gas turbine engine component cooling circuit

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EP1136652A1 (fr) * 2000-03-23 2001-09-26 General Electric Company Segment d'ailette de guidage à turbine avec circulation interne de refroidissement
US7695245B1 (en) * 2007-03-06 2010-04-13 Florida Turbine Technologies, Inc. Turbine airfoil with a multi-impingement cooled spar and shell
EP2628901A1 (fr) * 2012-02-15 2013-08-21 Siemens Aktiengesellschaft Aube de turbine à gaz avec refroidissement par impact
US20140075947A1 (en) * 2012-09-18 2014-03-20 United Technologies Corporation Gas turbine engine component cooling circuit

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3597859B1 (fr) * 2018-07-13 2023-08-30 Honeywell International Inc. Aube de turbine avec système de refroidissement tolérant à la poussière

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