EP3460190A1 - Structures d'amélioration de transfert de chaleur sur des nervures en ligne d'une cavité de surface portante d'une turbine à gaz - Google Patents

Structures d'amélioration de transfert de chaleur sur des nervures en ligne d'une cavité de surface portante d'une turbine à gaz Download PDF

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Publication number
EP3460190A1
EP3460190A1 EP17192319.6A EP17192319A EP3460190A1 EP 3460190 A1 EP3460190 A1 EP 3460190A1 EP 17192319 A EP17192319 A EP 17192319A EP 3460190 A1 EP3460190 A1 EP 3460190A1
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EP
European Patent Office
Prior art keywords
ribs
aerofoil
gas turbine
primary
turbine component
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.)
Withdrawn
Application number
EP17192319.6A
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German (de)
English (en)
Inventor
Anthony Davis
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Siemens AG
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Siemens AG
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Filing date
Publication date
Application filed by Siemens AG filed Critical Siemens AG
Priority to EP17192319.6A priority Critical patent/EP3460190A1/fr
Publication of EP3460190A1 publication Critical patent/EP3460190A1/fr
Withdrawn legal-status Critical Current

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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
    • 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
    • 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
    • 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
    • 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
    • F05D2260/2212—Improvement of heat transfer by creating turbulence
    • 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
    • F05D2260/2214—Improvement of heat transfer by increasing the heat transfer surface
    • F05D2260/22141—Improvement of heat transfer by increasing the heat transfer surface using fins or ribs

Definitions

  • the present invention relates to gas turbine engines, and more particularly to gas turbine components having an aerofoil, such as blades or vanes, with in-line ribs.
  • Cooling of gas turbine components is a major challenge and an area of interest in turbine technology.
  • a common technique for cooling a turbine component having an aerofoil, i.e. blade or vane is to have one or more internal passages, referred to as cooling channels or cooling passages, formed within an aerofoil cavity defined by an interior surface of the aerofoil and to flow a cooling fluid, such as cooling air, through the aerofoil cavity or through the cooling channels formed within the aerofoil cavity thereby cooling the interior surface of the aerofoil, and consequently cooling the aerofoil of the blade/vane.
  • the cooling fluid enters the aerofoil, i.e. the cooling air flows into the aerofoil cavity, though an opening of the aerofoil cavity from one radial side of the aerofoil such as a platform side or a root side in case of blades or vanes, or from both sides such as a tip and a root side in case of vanes, and then flows along and in contact with the interior surface, often in a serpentine flow path defined within the aerofoil cavity, and exits the aerofoil cavity for film cooling through holes that open on an external or exterior surface of the aerofoil or exits through slots formed, generally on a trailing edge of the aerofoil.
  • the cooling air is flowed through an impingement tube positioned within the aerofoil cavity and exited out from the impingement tube via impingement holes of the impingement tube in form of impingement jets directed towards the interior surface of the aerofoil.
  • the cooling air after impinging on the interior surface of the aerofoil flows along and in contact with the interior surface, and exits the aerofoil cavity for film cooling through holes that open on the external surface of the aerofoil or exits through slots formed, generally at or near the trailing edge of the aerofoil.
  • chordwise extending ribs are formed on the interior surface, and agreeing with the direction of flow of the cooling air extend chordwise or spanwise within the interior surface.
  • the chordwise extending ribs extend from a leading edge side towards the trailing edge side of the aerofoil and direct the cooling air towards the trailing edge, when the impingement jets impact the interior surface at the leading edge side.
  • the chordwise extending ribs also assist in locating the impingement tube within the aerofoil cavity.
  • the spanwise extending ribs are often formed within the cooling channels and direct the cooling air radially inwards or radially outwards, with respect to a rotational axis of the gas turbine, within the aerofoil cavity i.e. from the tip end of the aerofoil towards the platform end of the aerofoil and vice versa. It may be noted that in both aforementioned schemes of ribs arrangement the ribs extend longitudinally along the direction of flow of the cooling air, and hence the aforementioned spanwise or chordwise extending ribs are generally referred to as in-line ribs signifying their arrangement with respect to the direction of flow of the cooling air i.e. being in-line with the direction of flow of the cooling air.
  • the aforementioned ribs are aligned longitudinally on the interior surface of the aerofoil along the intended flow direction of the cooling air.
  • the ribs besides directing the cooling air increase a surface area of the interior surface and thus result in increased heat transfer from the interior surface to the cooling air flowing over, i.e. flowing along and in contact with, the interior surface.
  • further enhancement of heat transfer and thus an increase in cooling of the aerofoil is advantageous. Therefore, a technique is desired to increase heat transfer from interior surface of the aerofoil and thus increase the cooling of the aerofoil.
  • the object of the present invention is to provide a technique for increasing heat transfer from interior surface of the aerofoil to the cooling air flowing over the interior surface and thus increasing cooling of the aerofoil.
  • the invention presents a gas turbine component having enhanced cooling features on a surface of the in-line ribs.
  • the gas turbine component hereinafter also referred to as the component, has an aerofoil having an aerofoil cavity defined by an interior surface of the aerofoil.
  • the component may be a blade or a vane of a gas turbine.
  • the aerofoil cavity is configured to be flowed through by cooling air.
  • the component includes a plurality of primary ribs arranged on the interior surface of the aerofoil and within the aerofoil cavity. The primary ribs are spaced apart and extend from the interior surface towards an inside of the aerofoil cavity.
  • each of the primary ribs comprises a plurality of secondary heat transfer structures.
  • the secondary heat transfer structures are positioned on a surface of the primary ribs i.e. the secondary heat transfer structures, hereinafter also referred to as the secondary structures, extend from the surface of the primary ribs and have a free-hanging end disposed in the aerofoil cavity.
  • a surface of the secondary structures being continuous with the surface of the primary ribs results in an increase of the surface area of the primary ribs consequently resulting in increase in heat transfer, and thereby in resultant cooling, associated with the primary ribs.
  • the secondary structures are the enhanced cooling features.
  • the phrase 'longitudinally' with reference to the primary ribs as used herein means along a longitudinal axis of the primary ribs i.e. along a long axis or a lengthwise disposed axis of the primary rib.
  • the longitudinal axis of the primary rib is generally parallel to the interior surface of the aerofoil and may follow a contour of the interior surface.
  • the phrase 'extend longitudinally along a flow direction of the cooling air' means the longitudinal axis of the primary rib is aligned parallel to the flow direction of the cooling air.
  • the primary ribs extend longitudinally between a tip end and a platform end of the aerofoil i.e. the longitudinal axis of the primary ribs extends between the tip end and the platform end i.e. extends from the tip end to the platform end.
  • the primary ribs extend radially, with respect to a rotation axis of the gas turbine.
  • the cooling air entering the aerofoil cavity hereinafter also referred to as the cavity, generally enters through the platform end of the aerofoil or in other words from a root section in case where the component is a blade of the gas turbine engine and flows within the aerofoil cavity towards the tip end of the aerofoil.
  • the cooling air enters from a root section of the vane and/or a tip section of the vane and flows within the aerofoil of the vane generally radially inwards or outwards, respectively.
  • the cooling air generally flows in a cooling channel having a serpentine path formed by one or more hairpin turns of the cooling channel.
  • the primary ribs extend longitudinally between one or more sections of the cooling channel between the hairpin turns.
  • the primary ribs extend longitudinally between a leading edge and a trailing edge of the aerofoil i.e. the longitudinal axis of the primary ribs extends between the leading edge and the trailing edge i.e. extends from a side of the leading edge to a side of the trailing edge.
  • Such ribs are generally used to locate an impingement tube within the aerofoil cavity. Impingement jets ejected from the impingement tube, particularly from impingement holes present in the impingement tube, are directed towards the interior wall of the cavity.
  • the cooling air is directed generally in a flow direction from the leading edge towards the trailing edge by the primary ribs that extend longitudinally between the leading edge and the trailing edge.
  • the flow may also be directed in a general direction towards the leading edge if appropriate.
  • the cooling air after flowing towards the trailing edge is ejected out of the aerofoil through holes or slots for example holes or slots at the trailing edge, holes or slots at the leading edge.
  • the secondary ribs also increase the turbulence in the cooling air.
  • the secondary heat transfer structures are secondary ribs.
  • Ribs are devices or features that emulate a rod or a bar i.e. a rigid strip shaped and having a cross-section that is semicircular, parabolic or polygonal.
  • the secondary ribs extend longitudinally along the flow direction of the cooling air i.e. a longitudinal axis of the secondary rib is parallel to the longitudinal axis of the primary rib.
  • the phrase 'longitudinally' with reference to the secondary ribs as used herein means along the longitudinal axis of the secondary ribs i.e. along a long axis or a lengthwise disposed axis of the secondary rib.
  • the longitudinal axis of the secondary rib for this embodiment is generally parallel to the interior surface of the aerofoil.
  • the secondary ribs extend longitudinally at an angle to the flow direction of the cooling air i.e. the longitudinal axis of the secondary rib is at an angle to the longitudinal axis of the primary rib.
  • the phrase 'longitudinally' with reference to the secondary ribs as used herein means along the longitudinal axis of the secondary ribs i.e. along a long axis or a lengthwise disposed axis of the secondary rib.
  • the longitudinal axis of the secondary rib for this embodiment is at the aforementioned angle to the interior surface of the aerofoil.
  • the angle is a right angle, or in other words the longitudinal axis of the secondary ribs is mutually perpendicular to the longitudinal axis of the primary rib on which the secondary ribs are positioned and to the interior surface of the aerofoil.
  • a maximum cross-sectional extent of the secondary ribs is between 50% and 100% of a thickness of the primary rib at which the secondary ribs are positioned, particularly about 75% of the thickness of the primary rib at which the secondary ribs are positioned.
  • a height of the secondary ribs is between one time and two and half times of the maximum cross-sectional extent of the secondary rib.
  • the secondary heat transfer structures are pin-fins.
  • a diameter of the pin-fins is between 50% and 100% of a thickness of the primary rib at which the pin-fins are positioned, particularly about 75% of the thickness of the primary rib at which the pin-fins are positioned, and a height of the pin-fins is between one time and two and half times of the diameter of the pin-fin.
  • the secondary heat transfer structures are pimples or dimples.
  • a maximum cross-sectional extent of the pimples or dimples is between 50% and 100% of a thickness of the primary rib at which the pimples or dimples are positioned, and particularly about 75% of the thickness of the primary rib at which the pimples or dimples are positioned.
  • FIG. 1 shows an exemplary embodiment of a gas turbine engine 10 in a sectional view in which a gas turbine component 1 (shown in FIGs 2 , 5 , 6 , 8 and 9 to 15 ) is incorporated.
  • the gas turbine engine 10, hereinafter referred to as the engine 10 comprises, in flow series, an inlet 12, a compressor or compressor section 14, a combustor section 16 and a turbine section 18 which are generally arranged in flow series and generally about and in the direction of a rotational axis 20.
  • the gas turbine engine 10 further comprises a shaft 22 which is rotatable about the rotational axis 20 and which extends longitudinally through the gas turbine engine 10.
  • the shaft 22 drivingly connects the turbine section 18 to the compressor section 14.
  • air 24 which is taken in through the air inlet 12 is compressed by the compressor section 14 and delivered to the combustion section or burner section 16.
  • the burner section 16 comprises a longitudinal axis 35, a burner plenum 26, one or more combustion chambers 28 and at least one burner 30 fixed to each combustion chamber 28.
  • the combustion chambers 28 and the burners 30 are located inside the burner plenum 26.
  • the compressed air passing through the compressor 14 enters a diffuser 32 and is discharged from the diffuser 32 into the burner plenum 26 from where a portion of the air enters the burner 30 and is mixed with a gaseous or liquid fuel.
  • the air/fuel mixture is then burned and the combustion gas 34 or working gas from the combustion is channelled through the combustion chamber 28 to the turbine section 18 via a transition duct 17.
  • An inner surface 55 of the transition duct 17 defines a part of the hot gas path.
  • This exemplary gas turbine engine 10 has a cannular combustor section arrangement 16, which is constituted by an annular array of combustor cans 19 each having the burner 30 and the combustion chamber 28, the transition duct 17 has a generally circular inlet that interfaces with the combustor chamber 28 and an outlet in the form of an annular segment.
  • An annular array of transition duct outlets form an annulus for channelling the combustion gases to the turbine section 18.
  • the turbine section 18 comprises a number of blade carrying discs 36 attached to the shaft 22.
  • two discs 36 each carry an annular array of turbine blades 38 are depicted, however, the number of blade carrying discs could be different, i.e. only one disc or more than two discs.
  • guiding vanes 40 which are fixed to a stator 42 of the gas turbine engine 10, are disposed between the stages of annular arrays of turbine blades 38. Between the exit of the combustion chamber 28 and the leading turbine blades 38 inlet guiding vanes 44 are provided that turn the flow of working gas onto the turbine blades 38.
  • the combustion gas 34 from the combustion chamber 28 enters the turbine section 18 and drives the turbine blades 38 which in turn rotate the shaft 22.
  • the guiding vanes 40, 44 hereinafter also referred to as the vanes 40,44, serve to optimise the angle of the combustion or working gas 34 on the turbine blades 38.
  • the turbine section 18 drives the compressor section 14.
  • the compressor section 14 comprises an axial series of vane stages 46 and rotor blade stages 48.
  • the rotor blade stages 48 comprise a rotor disc supporting an annular array of blades.
  • the compressor section 14 also comprises a casing 50 that surrounds the rotor stages and supports the vane stages 48.
  • the guide vane stages include an annular array of radially extending vanes that are mounted to the casing 50. The vanes are provided to present gas flow at an optimal angle for the blades at a given engine operational point.
  • Some of the guide vane stages have variable vanes, where the angle of the vanes, about their own longitudinal axis, can be adjusted for angle according to air flow characteristics that can occur at different engine operations conditions.
  • the casing 50 defines a radially outer surface 52 of the passage 56 of the compressor 14.
  • a radially inner surface 54 of the passage 56 is at least partly defined by a rotor drum 53 of the rotor which is partly defined by the annular array of blades 48.
  • the present technique is described with reference to the above exemplary turbine engine having a single shaft or spool connecting a single, multi-stage compressor and a single, one or more stage turbine.
  • the present technique is equally applicable to two or three shaft engines and which can be used for industrial, aero or marine applications.
  • the cannular combustor section arrangement 16 is also used for exemplary purposes and it should be appreciated that the present technique is equally applicable to annular type and can type combustion chambers.
  • the gas turbine component 1 of the present technique may be the blade 38 as shown in FIG 1 or the vane 40,44 as shown in FIG 1 . It may be noted that the present technique has been primarily explained in details hereinafter with respect to an embodiment of the turbine blade 38 as the gas turbine component 1 of the present technique, however, it must be appreciated that the present technique is equally applicable and implemented similarly with respect to the turbine vane 40,44 or any other turbomachine component having an aerofoil and being cooled by a cooling fluid flowing through an aerofoil cavity formed within the aerofoil.
  • FIGs 2, 3 and 5 are used hereinafter to describe an exemplary embodiment of a gas turbine engine component 1 of the present technique.
  • FIG 4 has been used to depict a conventional design and for explaining difference between the conventional design as shown in FIG 4 and the gas turbine engine component 1 of the present technique as shown in FIG 5 .
  • the gas turbine component 1, hereinafter also referred to as the component 1, has an aerofoil 60.
  • the aerofoil 60 may extend from a platform (not shown in FIG 2 ) in the radial direction. From the platform also emanates a root (not shown in FIG 2 ) or a fixing part generally when the component 1 is embodied as the blade 38 (shown in FIG 1 ). The root or the fixing part may be used to attach the component 1 to the turbine disc 38 (shown in FIG 1 ). It may be noted that in some other embodiments of the component 1, for example when the component 1 is the vane 40,44 the aerofoil 60 may be affixed to fixed structures such as the stator discs or the housing 50, as shown in FIG 1 .
  • the aerofoil 60 includes a pressure side 67 and a suction side 68.
  • the pressure side 67 and the suction side 68 meet at a trailing edge 66 on one end and a leading edge 64 on another end.
  • the aerofoil 60 has a platform end 63 at which the aerofoil 60 is attached to the platform, and a tip end 61, radially outwards of the platform end 63.
  • the aerofoil 60 may be connected to a shroud (not shown in FIG 2 ) at the tip end 61 of the aerofoil 60. In some other embodiments the aerofoil 60 may be connected to a tip platform (not shown) instead of the shroud.
  • the interior surface 62 of the aerofoil 60 binds or defines an aerofoil cavity 65 formed within the aerofoil 60.
  • the aerofoil cavity 65 hereinafter also referred to as the cavity 65, extends radially or spanwise between the tip end 61 and the platform end 63 of the aerofoil 60.
  • an impingement tube 2 may be located within the cavity 65 .
  • a region of the aerofoil 60 marked with reference character A in FIG 2 has been shown in details in FIG 3 .
  • To cool the interior surface 62 of the aerofoil 60 cooling air is made to flow through the impingement tube 2 located within the cavity 65 and in vicinity of the interior surface 62.
  • the cooling air exits the impingement tube 2 via impingement holes 3 of the impingement tube 2 in form of impingement jets 4 directed towards the interior surface 62 of the aerofoil 60.
  • the cooling air is guided by ribs 70, referred to as primary ribs 60, formed on the interior surface 62 of the aerofoil 60.
  • the cooling air is guided along the interior surface 62 in a direction extending from the leading edge 64 towards the trailing edge 66.
  • the direction of flow of cooling air has been depicted in FIGs 2 and 3 by arrows marked with reference numeral 9.
  • the cooling air is directed towards the leading edge 64.
  • a plurality of the primary ribs 70 are arranged on the interior surface 62 of the aerofoil 60.
  • the primary ribs 70 are spaced apart, generally radially, from each other and extend from the interior surface 62 towards an inside of the aerofoil cavity 65.
  • the primary ribs 70 extend longitudinally along the flow direction 9 of the cooling air. In the example of FIGs 2, 3 and 5 the primary ribs 70 thus extend in the direction extending from the leading edge 64 towards the trailing edge 66.
  • the phrase 'longitudinally' with reference to the primary ribs 70 as used herein means along a longitudinal axis 71 of the primary ribs 70 i.e. along a long axis 71 or a lengthwise disposed axis 71 of the primary rib 70.
  • the longitudinal axis 71 of each of the primary ribs 70 is generally parallel to the interior surface 62 of the aerofoil 60 and may follow a contour of the interior surface 62.
  • the longitudinal axis 71 of each of the primary ribs 70 is parallelly aligned with the flow direction 9 of the cooling air as shown in FIG 3 .
  • the primary ribs 70 are thus chordwise extending i.e.
  • the longitudinal axis 71 of the primary ribs 70 extends chordwise or between the leading edge 64 and the trailing edge 66 of the aerofoil 60.
  • each of the primary ribs 70 comprises a plurality of secondary heat transfer structures 80.
  • the secondary heat transfer structures 80 are positioned on a surface 72 of the primary ribs 70 i.e. the secondary heat transfer structures 80, hereinafter also referred to as the secondary structures 80, extend from the surface 72 of the primary ribs 70 and have a free-hanging end disposed in the aerofoil cavity 65.
  • the secondary structures 80 are devices or features or elements that result in increase in a surface area of the surface 72 of the primary ribs 70.
  • the secondary structures 80 may be rib shaped, and therefore referred to as secondary ribs 82 as shown in FIG 5 , or may have other shapes as explained later with reference to FIGs 14 and 15 .
  • FIG 4 presents conventionally known chordwise extending ribs 70.
  • the conventionally known chordwise extending ribs 70 do not have the secondary structures 80 on their surface, and thus the surface 72 of the conventionally known chordwise extending ribs 70 is flat or smooth as shown in FIG 4 .
  • the surface 72 of the primary ribs 70 also extending chordwise, has the secondary structures 80 formed on the surface 72 of the primary ribs 70.
  • FIGs 6 and 8 are used hereinafter to describe another exemplary embodiment of the gas turbine engine component 1 of the present technique.
  • FIG 7 has been used to depict a conventional design and for explaining difference between the conventional design as shown in FIG 7 and the gas turbine engine component 1 of the present technique as shown in FIG 8 .
  • the component 1 has the aerofoil 60 extending from a platform 7 in the radial direction.
  • the platform end 63 of the aerofoil 60 is the part of the aerofoil 60 at which the aerofoil 60 is attached to the platform 7.
  • the tip end 61 is the radially outward end of the aerofoil 60, and opposite to the platform end 63. From the platform 7 also emanates the root 8 generally when the component 1 is embodied as the blade 38 (shown in FIG 1 ).
  • the interior surface 62 defines the aerofoil cavity 65.
  • the aerofoil cavity 65 may be designed such that, by forming cooling channels within the cavity 65, the cooling air enters the aerofoil cavity 65 from the platform end 63 and flows within the cavity 65 in a serpentine path as depicted by the arrows marked with reference numeral 9 in FIG 6 .
  • the cooling channel thereafter makes a hairpin turn and thus the cooling air flows within the cavity 65 from the tip end 61 towards the platform end 63.
  • the cooling channel may make further hairpin turns thus making the cooling air flow in the serpentine path.
  • the cooling channel depicted in FIG 6 is for exemplary purposes only. It may be appreciated by one skilled in the art that the cooling channel may have more tortuous path than the one depicted in FIG 6 . It may also be appreciated by one skilled in the art that the cooling channel may have less tortuous path than the one depicted in FIG 6 , for example the cooling path may be the cavity 65 without any hairpin bends.
  • FIG 8 A region of the aerofoil 60 marked with reference character B in FIG 6 has been shown in details in FIG 8 .
  • the primary ribs 70 are arranged on the interior surface 62 of the aerofoil 60.
  • the primary ribs 70 are spaced apart, chordwise, from each other and extend from the interior surface 62 towards the inside of the aerofoil cavity 65.
  • the primary ribs 70 extend longitudinally along the flow direction 9 of the cooling air. In the example of FIGs 6 and 8 the primary ribs 70 thus extend in the direction extending between the tip end 61 and the platform end 63 of the aerofoil 60.
  • the longitudinal axis 71 of the primary ribs 70 is aligned extending between the tip end 61 and the platform end 63 of the aerofoil 60 and is generally parallel to the interior surface 62 of the aerofoil 60, and may follow a contour of the interior surface 62. In other words, the longitudinal axis 71 of each of the primary ribs 70 is parallelly aligned with the flow direction 9 of the cooling air as shown in FIGs 6 and 8 .
  • the primary ribs 70 are thus spanwise extending i.e. extending longitudinally between the tip end 61 and the platform end 63 of the aerofoil 60, or in other words the longitudinal axis 71 of the primary ribs 70 extends spanwise.
  • each of the primary ribs 70 comprises the secondary structures 80.
  • the secondary structures 80 are positioned on the surface 72 of the primary ribs 70 i.e. the secondary structures extend from the surface 72 of the primary ribs 70 and have a free-hanging end disposed in the aerofoil cavity 65.
  • the secondary structures 80 are devices or features or elements that result in increase in a surface area of the surface 72 of the primary ribs 70.
  • the secondary structures 80 may be rib shaped, and therefore referred to as the secondary ribs 82 as shown in FIG 8 , or may have other shapes as explained later with reference to FIGs 14 and 15 .
  • FIG 7 presents conventionally known spanwise extending ribs 70.
  • the conventionally known spanwise extending ribs 70 do not have the secondary structures 80 on their surface, and thus the surface 72 of the conventionally known spanwise extending ribs 70 is flat or smooth as shown in FIG 7 .
  • the surface 72 of the primary ribs 70 also extending spanwise, has the secondary structures 80 formed on the surface 72 of the primary ribs 70.
  • the secondary structures 80 help in increasing the surface area for heat transfer between the primary ribs 70 and the cooling air. Additionally, the secondary structures 80 also help in inducing turbulence in the cooling air flow 9.
  • FIGs 9 and 10 show the cooling air flowing between two of the primary ribs 70.
  • FIG 9 shows the secondary structures 80 having rectangular cross-sections as for the secondary ribs 82 of FIGs 5 and 8 .
  • FIG 10 shows the secondary structures 80 having semi-circular cross-sections instead of the rectangular cross-section shown in FIGs 5 and 8 .
  • the cooling air flow 9 enters non-turbulently, i.e. emulating laminar flow, but due to the secondary structures 80 present on the surface 72 of the primary ribs 70 turbulence 5 is generated in the cooling air.
  • the secondary ribs 82 are formed on the surface 72 of the primary ribs 70.
  • the secondary ribs 82 are devices or features that emulate a rod or a bar i.e. a rigid strip shaped and having a cross-section that is semicircular, parabolic or polygonal.
  • FIG 12 shows three exemplary shapes of the secondary ribs 82 - a secondary rib 82a having a rectangular cross-section, another secondary rib 82b having a semi-circular or parabolic cross-section, and yet another secondary rib 82c having a triangular cross-section. It may be noted that further shapes of the secondary ribs 82 are also within the scope of the present technique, for example a pentagonal or hexagonal cross-section of the secondary ribs 82.
  • the secondary ribs 82 extend longitudinally along the flow direction 9 of the cooling air i.e. a longitudinal axis 81 of the secondary rib 82 is parallel to the longitudinal axis 71 of the primary rib 70.
  • the phrase 'longitudinally' with reference to the secondary ribs 82 as used herein means along the longitudinal axis 81 of the secondary ribs 82 i.e. along a long axis 81 or a lengthwise disposed axis 81 of the secondary rib 82.
  • the longitudinal axis 81 of the secondary rib 82 for this embodiment is generally parallel to the interior surface 62 of the aerofoil 60.
  • the secondary ribs 82 extend longitudinally at an angle 75 to the flow direction 9 of the cooling air i.e. the longitudinal axis 81 of the secondary rib 82 is at the angle 75 to the longitudinal axis 71 of the primary rib 70.
  • the phrase 'longitudinally' with reference to the secondary ribs 82 as used herein means along the longitudinal axis 81 of the secondary rib 82 i.e. along the long axis 81 or the lengthwise disposed axis 81 of the secondary rib 82.
  • the longitudinal axis 81 of the secondary rib 82 for this embodiment is at the aforementioned angle 75 to the interior surface 62 of the aerofoil 60.
  • the angle 75 may be 10 degree to 90 degree.
  • the angle 75 is a right angle, or in other words the longitudinal axis 81 of each of the secondary ribs 82 is mutually perpendicular to the longitudinal axis 71 of the primary rib 70 on which the secondary ribs 82 are positioned and to the interior surface 62 of the aerofoil 60.
  • the heat transfer to the cooling air from the secondary rib 82 varies with change in the angle 75, i.e. for example measure of heat transfer expressed via Nusselt number is smaller when the secondary ribs 82 are at right angle to the flow direction 9 of the cooling air compared to the heat transfer when the secondary ribs 82 are at another angle, for example 45 degree to the flow direction 9 of the cooling air.
  • the angle 75 is the smaller angle of the two complementary angles formed by the secondary rib 82 with the flow direction 9 of the cooling air, or in other words with the longitudinal axis 71 of the primary rib 70.
  • the increase in heat transfer also results in increase in pressure loss in the cooling air. So depending on the level of heat transfer required, and the pressure available, the choice of the angle 75 for orienting the secondary ribs 82 on the surface 72 of the primary rib 70 can be made.
  • a maximum cross-sectional extent L of the secondary ribs 82 is between 50% and 100% of a thickness T of the primary rib 70 at which the secondary ribs 82 are positioned, particularly about 75% of the thickness T of the primary rib 70 at which the secondary ribs 82 are positioned.
  • the maximum cross-sectional extent L for a given secondary rib 82 means greatest or maximum measure of length, for the given secondary rib 82, performed parallel to the surface 72 of the primary rib 70 and perpendicular to the longitudinal axis 81 of the secondary rib 82.
  • the maximum cross-sectional extent L for a given secondary rib 82 means a measure of length, for the given secondary rib 82, performed parallel to the surface 72 of the primary rib 70 and perpendicular to the longitudinal axis 81 of the secondary rib 82.
  • the thickness T of the primary rib is measured perpendicular to the surface 72 of the primary rib 70.
  • a height H of each of the secondary ribs 82 is between one time, i.e. equal to, and two and half times (2.5 times) of the maximum cross-sectional extent L of the secondary rib 82.
  • the height H is measured perpendicular to the surface 72 of the primary rib 70.
  • the secondary structures 80 are pin-fins 90.
  • a diameter d of the pin-fins 90 is between 50% and 100% of the thickness T, as aforementioned, of the primary rib 70 at which the pin-fins 90 are positioned, and particularly about 75% of the thickness T of the primary rib 70 at which the pin-fins 90 are positioned.
  • a height h of the pin-fins 90 is between one time, i.e. equal to, and two and half times (2.5 times) of the diameter d of the pin-fin 90.
  • the secondary structures 80 are pimples 95 or dimples 95.
  • a maximum cross-sectional extent L' of the pimples 95 or the dimples 95 is between 50% and 100% of the thickness T of the primary rib 70 at which the pimples 95 or dimples 95 are positioned, and particularly about 75% of the thickness T of the primary rib 70 at which the pimples 95 or the dimples 95 are positioned.
  • the maximum cross-sectional extent L' of the pimples 95 or the dimples 95 is the measure of a diameter of the pimple 95 or the dimple 95 performed at a level of the surface 72 of the primary rib 70.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP17192319.6A 2017-09-21 2017-09-21 Structures d'amélioration de transfert de chaleur sur des nervures en ligne d'une cavité de surface portante d'une turbine à gaz Withdrawn EP3460190A1 (fr)

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EP17192319.6A EP3460190A1 (fr) 2017-09-21 2017-09-21 Structures d'amélioration de transfert de chaleur sur des nervures en ligne d'une cavité de surface portante d'une turbine à gaz

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EP17192319.6A EP3460190A1 (fr) 2017-09-21 2017-09-21 Structures d'amélioration de transfert de chaleur sur des nervures en ligne d'une cavité de surface portante d'une turbine à gaz

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US11118462B2 (en) 2019-01-24 2021-09-14 Pratt & Whitney Canada Corp. Blade tip pocket rib
US11371359B2 (en) 2020-11-26 2022-06-28 Pratt & Whitney Canada Corp. Turbine blade for a gas turbine engine
CN114893254A (zh) * 2022-04-22 2022-08-12 中国联合重型燃气轮机技术有限公司 发动机叶片和燃气轮机
CN115875084A (zh) * 2023-03-02 2023-03-31 中国航发四川燃气涡轮研究院 应用于涡轮叶片压力面的层板冷却结构
JP7548778B2 (ja) 2019-11-06 2024-09-10 マン・エナジー・ソリューションズ・エスイー 衝突冷却によってガスタービン／ターボ機械の構成部品を冷却するための装置

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US20160069194A1 (en) * 2014-09-09 2016-03-10 Honeywell International Inc. Turbine blades and methods of forming turbine blades having lifted rib turbulator structures
US20170030202A1 (en) * 2015-07-29 2017-02-02 General Electric Company Article, airfoil component and method for forming article

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US5468125A (en) * 1994-12-20 1995-11-21 Alliedsignal Inc. Turbine blade with improved heat transfer surface
US6290462B1 (en) * 1998-03-26 2001-09-18 Mitsubishi Heavy Industries, Ltd. Gas turbine cooled blade
US20030049127A1 (en) * 2000-03-22 2003-03-13 Peter Tiemann Cooling system for a turbine blade
US20160069194A1 (en) * 2014-09-09 2016-03-10 Honeywell International Inc. Turbine blades and methods of forming turbine blades having lifted rib turbulator structures
US20170030202A1 (en) * 2015-07-29 2017-02-02 General Electric Company Article, airfoil component and method for forming article

Cited By (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US11118462B2 (en) 2019-01-24 2021-09-14 Pratt & Whitney Canada Corp. Blade tip pocket rib
JP7548778B2 (ja) 2019-11-06 2024-09-10 マン・エナジー・ソリューションズ・エスイー 衝突冷却によってガスタービン／ターボ機械の構成部品を冷却するための装置
US11371359B2 (en) 2020-11-26 2022-06-28 Pratt & Whitney Canada Corp. Turbine blade for a gas turbine engine
CN114893254A (zh) * 2022-04-22 2022-08-12 中国联合重型燃气轮机技术有限公司 发动机叶片和燃气轮机
CN115875084A (zh) * 2023-03-02 2023-03-31 中国航发四川燃气涡轮研究院 应用于涡轮叶片压力面的层板冷却结构
CN115875084B (zh) * 2023-03-02 2023-06-30 中国航发四川燃气涡轮研究院 应用于涡轮叶片压力面的层板冷却结构

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