WO2015195088A1 - Système de refroidissement d'un profil de turbine comprenant un système de refroidissement par impact d'un bord d'attaque - Google Patents

Système de refroidissement d'un profil de turbine comprenant un système de refroidissement par impact d'un bord d'attaque Download PDF

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Publication number
WO2015195088A1
WO2015195088A1 PCT/US2014/042615 US2014042615W WO2015195088A1 WO 2015195088 A1 WO2015195088 A1 WO 2015195088A1 US 2014042615 W US2014042615 W US 2014042615W WO 2015195088 A1 WO2015195088 A1 WO 2015195088A1
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WO
WIPO (PCT)
Prior art keywords
leading edge
wall
airfoil
cooling
pressure side
Prior art date
Application number
PCT/US2014/042615
Other languages
English (en)
Inventor
Jan H. Marsh
Ralph W. Matthews
Original Assignee
Siemens Energy, Inc.
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 Energy, Inc. filed Critical Siemens Energy, Inc.
Priority to PCT/US2014/042615 priority Critical patent/WO2015195088A1/fr
Publication of WO2015195088A1 publication Critical patent/WO2015195088A1/fr

Links

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
    • F01D5/186—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
    • F05D2240/00—Components
    • F05D2240/20—Rotors
    • F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
    • F05D2240/303—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor related to the leading edge of a rotor blade
    • 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

Definitions

  • This invention is directed generally to turbine airfoils, and more particularly to leading edge cooling systems in hollow turbine airfoils of gas turbine engines.
  • gas turbine engines typically include a compressor for compressing air, a combustor for mixing the compressed air with fuel and igniting the mixture, and a turbine blade assembly for producing power.
  • Combustors often operate at high temperatures that may exceed 2,500 degrees Fahrenheit.
  • Typical turbine combustor configurations expose turbine blade assemblies to these high temperatures.
  • turbine blades must be made of materials capable of withstanding such high temperatures.
  • turbine blades often contain cooling systems for prolonging the life of the blades and reducing the likelihood of failure as a result of excessive temperatures.
  • turbine blades are formed from a root portion having a platform at one end and an elongated portion forming a blade that extends outwardly from the platform coupled to the root portion.
  • the blade is ordinarily composed of a tip opposite the root section, a leading edge, and a trailing edge.
  • the inner aspects of most turbine blades typically contain an intricate maze of cooling channels forming a cooling system.
  • the cooling channels in a blade receive air from the compressor of the turbine engine and pass the air through the blade.
  • the cooling channels often include multiple flow paths that are designed to maintain all aspects of the turbine blade at a relatively uniform temperature. However, centrifugal forces and air flow at boundary layers often prevent some areas of the turbine blade from being
  • the leading edge of turbine airfoils includes a plurality of film cooling holes forming a showerhead. While the showerhead of film cooling holes cools the leading edge, conventional showerhead configurations are often inefficient.
  • a turbine airfoil usable in a turbine engine and having an internal cooling system with a leading edge impingement channel for enhanced cooling of the leading edge of the turbine airfoil without a leading edge film cooling showerhead is disclosed.
  • the internal cooling system may include an leading edge cooling supply channel formed from a leading edge wall having a leading edge tip that is advanced closer to an inner surface of the leading edge of the generally elongated, hollow airfoil than other aspects of the leading edge cooling supply channel.
  • the leading edge cooling supply channel may include one or more leading edge impingement orifices for directing cooling fluids to impinge on the inner surface of the leading edge of the airfoil in the leading edge impingement channel.
  • the impingement cooling fluid may be exhausted through one or more pressure side and suction side film cooling gill holes in the leading edge impingement channel that are aft of the leading edge of the airfoil.
  • the turbine airfoil may be a turbine blade or vane.
  • the turbine airfoil may be formed from a generally elongated, hollow airfoil having a leading edge, a trailing edge, a pressure side, a suction side, a tip at a first end, a root coupled to the airfoil at a second end generally opposite to the first end for supporting the airfoil and for coupling the airfoil to a disc, and an internal cooling system formed from at least one cavity in the elongated, hollow airfoil.
  • the internal cooling system may include a leading edge cooling supply channel and a leading edge impingement channel positioned within the generally elongated, hollow airfoil along the leading edge of the generally elongated, hollow airfoil.
  • the leading edge cooling supply channel may be formed from a leading edge wall having a leading edge tip that is advanced closer to an inner surface of the leading edge of the generally elongated, hollow airfoil than other aspects of the leading edge cooling supply channel. In such a position, the leading edge cooling supply channel may send impingement fluids against the inner surface of the leading edge to cool the leading edge.
  • the internal cooling system may also include one or more leading edge impingement orifices in the leading edge tip of the leading edge cooling supply channel for exhausting cooling fluids to impinge on the inner surface of the leading edge of the generally elongated, hollow airfoil in a leading edge impingement channel.
  • the leading edge impingement orifice in the leading edge tip of the leading edge cooling supply channel may be formed from a plurality of leading edge impingement orifices aligned into one or more spanwise extending rows of leading edge impingement orifices.
  • the leading edge cooling supply channel may be formed from a first leading edge wall that extends spanwise and a second leading edge wall that extends spanwise and is coupled to the first leading edge wall forming the leading edge tip, whereby the first leading edge wall is nonorthogonal to the second leading edge wall.
  • the first and second leading edge walls of the leading edge cooling supply channel may define a portion of the leading edge impingement channel formed from a pressure side section and a suction side section that is nonorthogonal to the pressure side section.
  • the pressure side section and the suction side section of the leading edge impingement channel may form a c- shaped cross-sectional leading edge impingement channel.
  • the first leading edge wall may also be aligned with an outer wall forming the pressure side of the generally elongated hollow airfoil, and the second leading edge wall may be aligned with an outer wall forming the suction side of the generally elongated hollow airfoil.
  • the leading edge cooling supply channel may be formed from a first aft edge wall that extends spanwise and a second aft edge wall that extends spanwise and is coupled to the first aft edge wall forming an aft edge tip, wherein the first aft edge wall is nonorthogonal to the second aft edge wall.
  • the internal cooling system may include one or more pressure side film cooling gill holes extending from the pressure side section through an outer wall forming the pressure side of the generally elongated hollow airfoil.
  • the pressure side film cooling gill hole extending from the pressure side section through the outer wall forming the pressure side of the generally elongated hollow airfoil may be formed from a plurality of pressure side film cooling gill holes forming a spanwise extending linear row of pressure side film cooling gill holes.
  • the internal cooling system may also include one or more suction side film cooling gill holes extending from the suction side section through an outer wall forming the suction side of the generally elongated hollow airfoil.
  • the suction side film cooling gill hole extending from the suction side section through the outer wall forming the suction side of the generally elongated hollow airfoil may be formed from a plurality of suction side film cooling gill holes forming a spanwise extending linear row of suction side film cooling gill holes.
  • the leading edge cooling supply channel may be offset from an outer wall forming the pressure side and an outer wall forming the suction side of the generally elongated hollow airfoil.
  • the leading edge cooling supply channel may be supported by a pressure side rib extending between the outer wall forming the pressure side of the generally elongated hollow airfoil and the first leading edge wall of the leading edge cooling supply channel and may be supported by a suction side rib extending between the outer wall forming the suction side of the generally elongated hollow airfoil and the second leading edge wall of the leading edge cooling supply channel.
  • leading edge cooling supply channel with leading edge impingement orifices provides enhanced near wall impingement with higher heat transfer augmentation without film cooling at the leading edge of the airfoil, unlike most conventional systems.
  • leading edge cooling supply channel provides an intermediate channel to feed near wall impingement at the leading edge.
  • leading edge cooling supply channel provides an intermediate channel to feed near wall impingement at the leading edge.
  • pressure side and suction side film cooling gill holes eject post impingement air away from the leading edge.
  • leading edge cooling supply channel Another advantage of the leading edge cooling supply channel is that the ability to reduce the distance between the leading edge cooling supply channel and the leading edge is that there is more flexibility in designing interior aspects of the airfoil, such as enabling ribs to be widened within the airfoil and for internal passages to be added.
  • Still another advantage of the internal cooling system is that a better cooling distribution may be achieved through the combination of impingement array and film cooling gill holes that pull flow towards the edges of the impingement passage along the outer wall.
  • Another advantage of the internal cooling system is that the internal cooling system experiences higher back side convective cooling.
  • Yet another advantage of the internal cooling system is that there is no need, or a reduced need, for film cooling and thus, all showerhead holes have been removed leaving only film cooling gill holes.
  • Another advantage of the internal cooling system is that the turbine airfoil has increased resistance to thermal barrier coating spallation because of the lack of leading edge showerhead.
  • Still another advantage of the internal cooling system is that internal cooling system experiences increased component cooling efficiency.
  • Another advantage of the internal cooling system is that the internal cooling system experiences a reduction in cooling fluid flow due to less film cooling requirements as well as improved back side cooling due to better distribution and higher magnitude of cold side heat transfer coefficients.
  • Figure 1 is a perspective view of a turbine airfoil having features according to the invention.
  • Figure 2 is a cross-sectional view of the turbine airfoil shown in Figure 1 taken along section line 2-2 in Figure 1 .
  • Figure 3 is a schematic diagram of the internal cooling system within the turbine airfoil of Figure 2.
  • Figure 4 is a detail view of the leading edge impingement channel and the leading edge cooling supply channel shown at detail line 4-4 in Figure 2.
  • a turbine airfoil 10 usable in a turbine engine and having an internal cooling system 14 with a leading edge impingement channel 16 for enhanced cooling of the leading edge 18 of the turbine airfoil 10 without a leading edge film cooling showerhead is disclosed.
  • the internal cooling system 14 may include an leading edge cooling supply channel 20 formed from a leading edge wall 22 having a leading edge tip 24 that is advanced closer to an inner surface 26 of the leading edge 18 of the generally elongated, hollow airfoil 28 than other aspects of the leading edge cooling supply channel 20.
  • the leading edge cooling supply channel 20 may include one or more leading edge impingement orifices 30 for directing cooling fluids to impinge on the inner surface 26 of the leading edge 18 of the airfoil 28 in the leading edge impingement channel 20.
  • the impingement cooling fluid may be exhausted through one or more pressure side and suction side film cooling gill holes 32, 34 in the leading edge impingement channel 20 that are aft of the leading edge 18 of the airfoil 10.
  • the turbine airfoil 10 may be a turbine blade or vane.
  • the turbine airfoil 10 may be formed from a generally elongated, hollow airfoil 28 having a leading edge 18, a trailing edge 36, a pressure side 48, a suction side 50, a tip 38 at a first end 40, a root 42 coupled to the airfoil 10 at a second end 44 generally opposite to the first end 40 for supporting the airfoil 10 and for coupling the airfoil 10 to a disc, and a cooling system 14 formed from at least one cavity 46 in the elongated, hollow airfoil 28.
  • the internal cooling system 14 may include a leading edge cooling supply channel 20 and a leading edge impingement channel 16 positioned within the generally elongated, hollow airfoil 28 along the leading edge 18 of the generally elongated, hollow airfoil 28, as shown in Figures 2- 4.
  • the leading edge cooling supply channel 20 may be formed from a leading edge wall 22 having a leading edge tip 24 that is advanced closer to the inner surface 26 of the leading edge 18 of the generally elongated, hollow airfoil 28 than other aspects of the leading edge cooling supply channel 20.
  • the leading edge cooling supply channel 20 may be formed from a first leading edge wall 52 that extends spanwise and a second leading edge wall 54 that extends spanwise and is coupled to the first leading edge wall 52 forming the leading edge tip 24.
  • the first leading edge wall 52 may be nonorthogonal to the second leading edge wall 54.
  • the first leading edge wall 52 may be aligned with an outer wall 56 forming the pressure side 48 of the generally elongated hollow airfoil 28.
  • the second leading edge wall 54 may be aligned with an outer wall 58 forming the suction side 50 of the generally elongated hollow airfoil 28.
  • the leading edge cooling supply channel 20 may be offset from the outer wall 56 forming the pressure side 48 and the outer wall 58 forming the suction side 50 of the generally elongated hollow airfoil 28.
  • the leading edge cooling supply channel 20 may be supported by a pressure side rib 60 extending between the outer wall 56 forming the pressure side 48 of the generally elongated hollow airfoil 28 and the first leading edge wall 52 of the leading edge cooling supply channel 20.
  • the leading edge cooling supply channel 20 may also be supported by a suction side rib 62 extending between the outer wall 58 forming the suction side 50 of the generally elongated hollow airfoil 28 and the second leading edge wall 54 of the leading edge cooling supply channel 20.
  • the leading edge cooling supply channel 20 may also be formed from a first aft edge wall 80 that extends spanwise and a second aft edge wall 82 that extends spanwise and is coupled to the first aft edge wall 80 forming an aft edge tip 84.
  • the first aft edge wall 80 may be nonorthogonal to the second aft edge wall 82.
  • the internal cooling system 14 may include one or more leading edge impingement orifices 30 in the leading edge tip 24 of the leading edge cooling supply channel 20 for exhausting cooling fluids to impinge on the inner surface 26 of the leading edge 18 of the generally elongated, hollow airfoil 28 in a leading edge impingement channel 16.
  • the internal cooling system 14 may include a plurality of leading edge impingement orifices 30 aligned into a spanwise extending row of leading edge impingement orifices 30.
  • leading edge impingement orifices 30 there may be multiple spanwise extending rows of leading edge impingement orifices 30, such as, but not limited to, a first spanwise extending row 64 at a stagnation line 66, a second spanwise extending row 68 on the pressure side 48 of the stagnation line 66 and a third spanwise extending row 70 on the suction side 50 of the stagnation line 66.
  • the leading edge impingement orifices 30 may have any appropriate sized opening and cross-sectional area and shape.
  • the first and second leading edge walls 52, 54 of the leading edge cooling supply channel 20 may define a portion of the leading edge impingement channel 16 formed from a pressure side section 72 and a suction side section 74 that is nonorthogonal to the pressure side section 72.
  • the pressure side section 72 and the suction side section 74 of the leading edge impingement channel 16 may form a c- shaped cross-sectional leading edge impingement channel 16.
  • the internal cooling system 14 may also include one or more pressure side film cooling gill holes 32 extending from the pressure side section 72 through the outer wall 56 forming the pressure side 48 of the generally elongated hollow airfoil 28.
  • the internal cooling system 14 may also include a plurality of pressure side film cooling gill holes 32 forming a spanwise extending linear row of pressure side film cooling gill holes 32.
  • the internal cooling system 14 may also include one or more suction side film cooling gill holes 34 extending from the suction side section 74 through the outer wall 58 forming the suction side 50 of the generally elongated hollow airfoil 28.
  • the internal cooling system 14 may also include a plurality of suction side film cooling gill holes 34 forming a spanwise extending linear row of suction side film cooling gill holes 34.
  • the inlets of the pressure and suction side film cooling gill holes 32, 34 may be aft of the leading edge tip 24 of the leading edge cooling supply channel 20.
  • the internal cooling system 14 may include the leading edge impingement orifices 30 but may not exhaust cooling fluids through the leading edge 18 of the airfoil 10. Rather, the cooling fluids may be exhausted through the radially inner or outer ends 40, 44 of the leading edge impingement channel 16. This configuration may develop significant cross flow near the tip 38 or elsewhere, which may degrade the effectiveness of the leading edge impingement orifices 30. However, the reduced distance between the leading edge tip 24 housing the leading edge impingement orifices 30 and the inner surface 26 of the leading edge 18 of the airfoil 10 should lower the negative impact versus conventional configurations.
  • cooling fluids such as, but not limited to, air
  • the cooling fluids may enter the leading edge cooling supply channel 20 and flow spanwise throughout the leading edge cooling supply channel 20.
  • the cooling fluids may flow through the leading edge
  • the cooling fluids may increase in temperature due to convection and may flow along the inner surface forming the pressure and suction sides 48, 50.
  • the cooling fluids may be exhausted from the leading edge impingement channel 16 through the pressure side and suction side film cooling gill holes 32, 34.

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

Abstract

L'invention porte sur un profil de turbine (10) pouvant être utilisé dans un moteur de turbine et doté d'un système interne de refroidissement (14) comprenant un canal d'impact (16) du bord d'attaque pour améliorer le refroidissement du bord d'attaque (18) du profil de turbine (10) sans l'utilisation d'une pomme d'arrosoir de refroidissement du film du bord d'attaque. Le système de refroidissement interne (14) peut comprendre un canal d'alimentation de refroidissement du bord d'attaque (20) formé à partir d'une paroi de bord d'attaque (22) et dont la pointe du bord d'attaque (24) est située plus proche d'une surface intérieure (26) du bord d'attaque (18) du profil creux (28) généralement allongé, que les autres parties du canal d'alimentation de refroidissement de bord d'attaque (20). Le canal d'alimentation de refroidissement du bord d'attaque (20) peut comprendre un ou plusieurs orifices d'impact du bord d'attaque (30) pour acheminer des fluides de refroidissement à entrer en collision avec la surface interne (26) du bord d'attaque (18) du profil (28). Le fluide de refroidissement par impact peut être évacué à travers un ou plusieurs trous de refroidissement par film (32, 34) situés du côté pression et du côté aspiration dans le canal d'impact du bord d'attaque (16).
PCT/US2014/042615 2014-06-17 2014-06-17 Système de refroidissement d'un profil de turbine comprenant un système de refroidissement par impact d'un bord d'attaque WO2015195088A1 (fr)

Priority Applications (1)

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PCT/US2014/042615 WO2015195088A1 (fr) 2014-06-17 2014-06-17 Système de refroidissement d'un profil de turbine comprenant un système de refroidissement par impact d'un bord d'attaque

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PCT/US2014/042615 WO2015195088A1 (fr) 2014-06-17 2014-06-17 Système de refroidissement d'un profil de turbine comprenant un système de refroidissement par impact d'un bord d'attaque

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Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2017074404A1 (fr) * 2015-10-30 2017-05-04 Siemens Aktiengesellschaft Profil aérodynamique de turbine avec refroidissement par impact décalé sur le bord de fuite
EP3550110A1 (fr) * 2018-04-04 2019-10-09 United Technologies Corporation Profil aérodynamique ayant un schéma de refroidissement de bord d'attaque à compensation de frappe arrière
US10577942B2 (en) 2016-11-17 2020-03-03 General Electric Company Double impingement slot cap assembly

Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE19617556A1 (de) * 1996-05-02 1997-11-06 Asea Brown Boveri Thermisch belastete Schaufel für eine Strömungsmaschine
EP1380724A2 (fr) * 2002-07-11 2004-01-14 Mitsubishi Heavy Industries, Ltd. Aube de turbine refroidie
US20130280091A1 (en) * 2012-04-24 2013-10-24 Mark F. Zelesky Gas turbine engine airfoil impingement cooling

Patent Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE19617556A1 (de) * 1996-05-02 1997-11-06 Asea Brown Boveri Thermisch belastete Schaufel für eine Strömungsmaschine
EP1380724A2 (fr) * 2002-07-11 2004-01-14 Mitsubishi Heavy Industries, Ltd. Aube de turbine refroidie
US20130280091A1 (en) * 2012-04-24 2013-10-24 Mark F. Zelesky Gas turbine engine airfoil impingement cooling

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2017074404A1 (fr) * 2015-10-30 2017-05-04 Siemens Aktiengesellschaft Profil aérodynamique de turbine avec refroidissement par impact décalé sur le bord de fuite
US10577942B2 (en) 2016-11-17 2020-03-03 General Electric Company Double impingement slot cap assembly
EP3550110A1 (fr) * 2018-04-04 2019-10-09 United Technologies Corporation Profil aérodynamique ayant un schéma de refroidissement de bord d'attaque à compensation de frappe arrière
US20190309631A1 (en) * 2018-04-04 2019-10-10 United Technologies Corporation Airfoil having leading edge cooling scheme with backstrike compensation

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