WO2015095253A1 - Turbine airfoil vane with an impingement insert having a plurality of impingement nozzles - Google Patents

Turbine airfoil vane with an impingement insert having a plurality of impingement nozzles Download PDF

Info

Publication number
WO2015095253A1
WO2015095253A1 PCT/US2014/070702 US2014070702W WO2015095253A1 WO 2015095253 A1 WO2015095253 A1 WO 2015095253A1 US 2014070702 W US2014070702 W US 2014070702W WO 2015095253 A1 WO2015095253 A1 WO 2015095253A1
Authority
WO
WIPO (PCT)
Prior art keywords
impingement
cooling
wall
orifice
turbine vane
Prior art date
Application number
PCT/US2014/070702
Other languages
French (fr)
Inventor
Ching-Pang Lee
Original Assignee
Siemens Aktiengesellschaft
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
Priority claimed from US14/133,773 external-priority patent/US9347324B2/en
Application filed by Siemens Aktiengesellschaft, Siemens Energy, Inc. filed Critical Siemens Aktiengesellschaft
Publication of WO2015095253A1 publication Critical patent/WO2015095253A1/en

Links

Classifications

    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01DNON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12Blades
    • F01D5/14Form or construction
    • F01D5/18Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/187Convection cooling
    • F01D5/188Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall
    • F01D5/189Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall the insert having a tubular cross-section, e.g. airfoil shape
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05DINDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00Components
    • F05D2240/10Stators
    • F05D2240/12Fluid guiding means, e.g. vanes
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05DINDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00Function
    • F05D2260/20Heat transfer, e.g. cooling
    • F05D2260/201Heat transfer, e.g. cooling by impingement of a fluid

Definitions

  • This invention is directed generally to turbine airfoil vanes, and more particularly to hollow turbine airfoil vanes having an impingement insert for passing fluids, such as air, to cool the airfoils.
  • 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 vane and blade assemblies to these high temperatures.
  • turbine vanes and blades must be made of materials capable of
  • turbine vanes and blades often contain cooling systems for prolonging the life of the vanes and blades and reducing the likelihood of failure as a result of excessive temperatures.
  • turbine vanes are formed from an elongated portion forming a vane having one end configured to be coupled to a vane carrier and an opposite end configured to be movably coupled to an inner endwall.
  • the vane is ordinarily composed of a leading edge, a trailing edge, a suction side, and a pressure side.
  • the inner aspects of most turbine vanes typically contain an intricate maze of cooling circuits forming a cooling system.
  • the cooling circuits in the vanes receive cooling fluid, e.g., air from the compressor of the turbine engine, and pass the fluid through the ends of the vane adapted to be coupled to the vane carrier.
  • the cooling circuits often include multiple flow paths that are designed to maintain all aspects of the turbine vane at a relatively uniform temperature. At least some of the fluid passing through these cooling circuits is exhausted through orifices in the leading edge, trailing edge, suction side, and pressure side of the vane.
  • the cooling system may include an impingement plate 3 with a plurality of impingement holes 4 for directing cooling fluids to impinge on the outer wall 6 forming a turbine airfoil.
  • the impingement plate 3 may be offset from the outer wall 6 a conventional distance.
  • the impingement plate 3 may be generally flat and reside in a single plane. In this configuration, the cross flow of cooling fluids often disrupts the impingement jets directed towards the outer wall, thereby negatively impacting the cooling function of the impingement jets. While advances have been made in the cooling systems in turbine vanes, a need still exists for a turbine vane having increased cooling efficiency for dissipating heat and passing a sufficient amount of cooling fluid through the vane.
  • a turbine vane comprising a generally elongated hollow airfoil and a cooling system.
  • the airfoil comprises an outer wall including a leading edge, a trailing edge, a pressure side, a suction side, an outer endwall at a first end, and an inner endwall at a second end opposite the first end.
  • the cooling system is positioned within the airfoil and comprises a cooling chamber and an impingement insert positioned in the cooling chamber. The impingement insert and an inner surface of the airfoil outer wall define a cooling channel therebetween.
  • the impingement insert includes a plurality of impingement nozzles extending toward the inner surface of the outer wall and a plurality of impingement orifices. At least one of the impingement orifices is arranged in a non-aligned pattern with respect to at least one adjacent impingement orifice such that cooling fluid passing out of the at least one impingement orifice does not directly flow into a centerline of a cooling fluid flowpath of cooling fluid passing out of the at least one adjacent
  • Each impingement orifice may be arranged in a non-aligned pattern with respect to at least one adjacent impingement orifice such that cooling fluid passing out of each impingement orifice does not directly flow into a centerline of a cooling fluid flowpath of cooling fluid passing out of the at least one adjacent impingement orifice.
  • the orifices may be arranged in a staggered pattern comprising alternating first and second rows that are offset from one another in a flow direction of cooling fluid through the cooling channel.
  • Each of the impingement orifices may be formed in an outermost aspect of a corresponding impingement nozzle, and the impingement nozzles may have a generally cylindrical cross sectional area.
  • Each impingement nozzle may include at least one impingement orifice for directing cooling fluids orthogonally away from the impingement insert.
  • a distance between the outermost aspect of the impingement nozzle and the inner surface of the outer wall may be less than half of a distance between an innermost aspect of the impingement insert and the inner surface of the outer wall.
  • Only select ones of the impingement nozzles may include a corresponding impingement orifice formed therein.
  • a turbine vane comprising a generally elongated hollow airfoil and a cooling system.
  • the airfoil comprises an outer wall including a leading edge, a trailing edge, a pressure side, a suction side, an outer endwall at a first end, and an inner endwall at a second end opposite the first end.
  • the cooling system is positioned within the airfoil and comprises a cooling chamber and an impingement insert positioned in the cooling chamber. The impingement insert and an inner surface of the airfoil outer wall define a cooling channel therebetween.
  • the impingement insert includes a plurality of impingement nozzles extending toward the inner surface of the outer wall and a plurality of impingement orifices.
  • the impingement orifices are arranged in a staggered pattern comprising alternating first and second rows that are offset from one another in a flow direction of cooling fluid through the cooling channel such that cooling fluid passing out of each respective impingement orifice does not directly flow into a centerline of a cooling fluid flowpath of cooling fluid passing out of the adjacent upstream impingement orifice.
  • a turbine vane comprising a generally elongated hollow airfoil and a cooling system.
  • the airfoil comprises an outer wall including a leading edge, a trailing edge, a pressure side, a suction side, an outer endwall at a first end, and an inner endwall at a second end opposite the first end.
  • the cooling system is positioned within the airfoil and comprises a cooling chamber and an impingement insert positioned in the cooling chamber.
  • the impingement insert and an inner surface of the airfoil outer wall define a cooling channel therebetween.
  • the impingement insert includes a plurality of impingement nozzles extending toward the inner surface of the outer wall and a plurality of impingement orifices.
  • the impingement orifices are formed in corresponding impingement nozzles and are arranged in a non-aligned pattern with respect to at least one adjacent impingement orifice such that cooling fluid passing out of the impingement orifices does not directly flow into a centerline of a cooling fluid flowpath of cooling fluid passing out of the at least one adjacent impingement orifice.
  • a distance between an outermost aspect of at least one impingement nozzle including an impingement orifice and the inner surface of the outer wall is less than half of a distance between an innermost aspect of the impingement insert and the inner surface of the outer wall.
  • FIG. 1 is a perspective view of a turbine vane having features according to the instant invention.
  • FIG. 2 is a cross-sectional view of the turbine vane shown in FIG. 1 taken along line 2-2.
  • FIG. 3 is a cross-sectional, detailed view taken of an outer wall of a conventional turbine airfoil with an impingement insert.
  • FIG. 4 is a cross-sectional, detailed view taken at detail line 4-4 in FIG. 2 displaying an impingement insert with a plurality of impingement nozzles.
  • FIG. 5 is a partial view of the inner surface of the outer wall taken along line 5-5 in FIG. 4 showing the impingement jets striking the outer wall and cross flow flowing therebetween.
  • FIGS. 6 and 7 are views similar to those of FIGS. 4 and 5 illustrating an impingement insert in accordance with another embodiment of the instant invention.
  • the turbine vane 10 may include one or more cooling systems 12 with an impingement plate 14, also referred to herein as an impingement insert 14, having one or more impingement nozzles 16.
  • the turbine vane impingement nozzles 16 may extend towards an outer wall 20 forming the turbine vane 10 and may reduce the mixing of cooling fluids with impingement jets 22. Instead, the nozzles 16 may terminate within close proximity of the outer wall 20, thereby reducing the effect of cooling fluid cross flow 62.
  • the cooling system 12 may be configured to cool internal and external aspects of the turbine vane 10 usable in a turbine engine.
  • the turbine airfoil cooling system 12 may be configured to be included within a stationary turbine vane 10, as shown in FIGS. 1 -4.
  • the cooling system 12 may include one or more cooling chambers 26.
  • the cooling chambers 26 may include one or more midcord cooling chambers 28 positioned in the outer wall 20.
  • the turbine vane 10 may be formed from a generally elongated hollow airfoil 30 having an outer surface 32 adapted for use, for example, in an axial flow turbine engine.
  • Outer surface 32 may have a generally concave shaped portion forming the pressure side 34 and a generally convex shaped portion forming the suction side 36.
  • the turbine vane 10 may also include an outer endwall 38 at a first end 40 adapted to be coupled to a hook attachment and may include an inner endwall 42 at a second end 44.
  • the airfoil 30 may also include a leading edge 46 and a trailing edge 48 opposite the leading edge 46.
  • the turbine vane 10 may include an impingement insert 14 positioned in internal aspects of a central cooling chamber 26 of the cooling system 12.
  • the impingement insert 14 may include a plurality of impingement nozzles 16, as shown in FIG. 4, extending from the impingement insert 14.
  • the nozzles 16 may extend toward an inner surface 52 of the outer wall 20 from the impingement insert 14.
  • One or more of the nozzles 16 may include one or more impingement orifices 54.
  • Each nozzle 16 may include at least one impingement orifice 54 positioned at an outermost aspect 56 of the nozzle 16 for directing cooling fluids orthogonally away from the impingement insert 14.
  • one or more impingement nozzles 16 may be generally cylindrical. As such, a plurality of impingement nozzles 16 may be generally cylindrical. In other embodiments, one or more impingement nozzles 16 may have a cross-sectional area formed as a cylinder, a rectangle, a triangle, a semicircle, and other appropriate shapes. The impingement nozzles 16 may also be configured with a conical shape such that a cross-sectional area at a base 58 is greater than a cross- sectional area at the outermost aspect 56. One or a plurality of impingement nozzles 16 may be configured have a generally conical shape and may include one or more impingement orifices 54.
  • the impingement nozzles 16 may be aligned into rows, as shown in FIG. 5 through depiction of the impingement jets 60.
  • the rows may extend in a generally spanwise direction, in a generally chordwise direction or other appropriate direction. Adjacent rows may be offset from each other as will be discussed in greater detail below.
  • the outermost aspect 56 of the impingement nozzle 16 and the impingement orifices 54 may be located a distance 64 that is less than a conventional distance 8 between a conventional impingement plate 3 with holes 4 and an outer wall 6 of a conventional vane, see FIGS. 3 and 4.
  • a distance between an innermost aspect 68 of the impingement insert 14 and the inner surface 52 of the outer wall 20 is greater than a conventional distance 8 between the conventional impingement plate 3 with holes 4 and the outer wall 6 of the conventional vane, and is greater than twice the distance 64 between the outermost aspect 56 of the impingement nozzle 16 and the inner surface 52 of the outer wall 20, i.e., the distance 64 between the outermost aspect 56 of the impingement nozzle 16 and the inner surface 52 of the outer wall 20 is less than half of the distance between the innermost aspect 68 of the impingement insert 14 and the inner surface 52 of the outer wall 20.
  • the cross- sectional areas between the impingement nozzles 16 and the outer wall 20 is less than the distance 64 between the impingement insert 14 and the outer wall 20.
  • the impingement nozzles 16 may be placed in closer position relative to outer wall 20 without changing overall volume of cooling fluid flow through a cooling channel 70 formed between the outer wall 20 and the impingement insert 14.
  • cooling fluids may flow from a cooling fluid supply source (not shown) into the cooling system 12.
  • the cooling fluids may be passed into the cooling channel 70 formed between the impingement insert 14 and the outer wall 20 through the impingement nozzles 16 and the impingement orifices 54 positioned at the outermost aspect 56 of the nozzles 16.
  • the size and configuration of the jet 60 of cooling fluids flowing from the nozzle 16 is controlled by the shape and size of the nozzle 16.
  • the impingement jets 60 may be generally circular when the cooling fluids strike the inner surface 52 of the outer wall 20, as shown in FIG. 5.
  • the cooling fluids After the cooling fluids impinge upon the inner surface 52 of the outer wall 20, the cooling fluids form a cross flow 62 flowing generally along the outer wall 20. Because the nozzles 16 extend to within close proximity of the outer wall 20, the impingement jets 60 have sufficient velocity such that the cross flow 62 does not disrupt the impingement jets 60. The cooling fluids flowing from the impingement nozzles 16 reduce the temperature of the outer wall 20.
  • the general structure of the airfoil outer wall 120 according to this aspect of the invention may be similar to that of the airfoil outer wall 20 as described above with reference to FIGS. 1 -5.
  • the impingement insert 1 14 is part of a cooling system 1 12 and includes a plurality of impingement nozzles 1 16 extending toward the inner surface 152 of the airfoil outer wall 120. Similar to the impingement nozzles 16 discussed above, the impingement nozzles 1 16 may have a generally cylindrical cross sectional area, as shown in FIG. 6.
  • Select ones or all of the impingement nozzles 1 16 include at least one impingement orifice 154 located at an outermost aspect 156 of the impingement nozzle 1 16 for directing cooling fluid CF orthogonally away from the impingement insert 1 14, see FIG. 6.
  • a distance Di between the outermost aspect 156 of the impingement nozzles 1 16 and the inner surface 152 of the airfoil outer wall 120 may be less than half of a distance D 2 between an innermost aspect 1 14A of the impingement insert 1 14 and the inner surface 152 of the outer wall 120.
  • cooling fluid CF is discharged from the impingement nozzles 1 1 6 generally close to the inner surface 1 52 of the outer wall 120 to maximize impingement cooling of the outer wall 120 provided by the cooling fluid CF and to reduce disruption of the post impingement cooling fluid CF flowing normal to the impinging jets 1 60 within the cooling channel 1 70, as will be discussed further below.
  • At least one of the impingement orifices 1 54 (and all of the impingement orifices 1 54 shown in Fig. 7) is arranged in a non-aligned pattern with respect to at least one adjacent impingement orifice 154.
  • the impingement orifices 1 54 are arranged in a staggered pattern comprising alternating first and second rows Ri , R 2 that are offset from one another in a flow direction F D of the post impingement cooling fluid CF flowing in parallel to the outer wall 120 through the cooling channel 1 70.
  • cooling fluid CF passing out of the at least one impingement orifice 1 54 does not directly flow into a centerline CCF of a cooling fluid flowpath FCF of cooling fluid CF passing out of the at least one adjacent impingement orifice 1 54.
  • impingement nozzles 1 1 6 may include a corresponding impingement orifice 1 54 formed therein.
  • Ones of the impingement nozzles 1 1 6 that do not include an impingement orifice could be provided to effect a more turbulent flow of cooling fluid CF through the cooling channel 1 70 to increase cooling provided to the outer wall 120 by the cooling fluid C F .
  • cooling fluid CF may flow from a cooling fluid supply source (not shown) into the cooling system 1 12.
  • the cooling fluid CF may be passed into the cooling channel 1 70 of the cooling system 1 12 between the impingement insert 1 14 and the outer wall 120 through the impingement nozzles 1 1 6 and the impingement orifices 1 54 positioned at the outermost aspect 1 56 of the nozzles 1 1 6.
  • the size and configuration of the jets 1 60 of the cooling fluid CF discharged from the nozzles 1 1 6 is controlled by the shape and size of the nozzles 1 1 6.
  • the impingement jets 1 60 may be generally circular when the cooling fluid CF strikes the inner surface 152 of the outer wall 120, as shown in FIG. 7. After the cooling fluid CF impinges upon the inner surface 152 of the outer wall 120, the cooling fluid CF flows through the cooling channel 170 along the cooling fluid flowpath FCF generally along the outer wall 120. Because the nozzles 1 16 extend to within close proximity of the outer wall 120, the impingement jets 160 have sufficient velocity such that the cooling fluid CF flowing along their respective cooling fluid flowpaths FCF does not disrupt the impingement jets 160.
  • the nozzles 1 16 of the present aspect of the invention since the cooling fluid CF passing out of the impingement orifices 154 does not directly flow into the centerlines CCF of the cooling fluid flowpaths FCF of the cooling fluid CF passing out of the adjacent impingement orifices 154.
  • the cooling fluid CF flowing from the impingement nozzles 1 16 reduces the temperature of the outer wall 120.

Landscapes

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

Abstract

A turbine vane includes a generally elongated hollow airfoil and a cooling system. The cooling system is positioned within the airfoil and includes a cooling chamber and an impingement insert positioned in the cooling chamber. The impingement insert and an inner surface of an outer wall of the airfoil define a cooling channel therebetween. The impingement insert includes a plurality of impingement nozzles extending toward the inner surface of the outer wall and a plurality of impingement orifices. At least one of the impingement orifices is arranged in a non-aligned pattern with respect to at least one adjacent impingement orifice such that cooling fluid passing out of the at least one impingement orifice does not directly flow into a centerline of a cooling fluid flowpath of cooling fluid passing out of the at least one adjacent impingement orifice.

Description

TURBINE AIRFOIL VANE WITH AN IMPINGEMENT INSERT HAVING A PLURALITY
OF IMPINGEMENT NOZZLES
CROSS REFERENCE TO RELATED APPLICATIONS
This application is a Continuation-ln-Part of U.S. Patent Application Serial No. 12/885,740, filed September 20, 2012, entitled "TURBINE AIRFOIL VANE WITH AN IMPINGEMENT INSERT HAVING A PLURALITY OF IMPINGEMENT NOZZLES" by Ching-Pang Lee, the entire disclosure of which is incorporated by reference herein.
FIELD OF THE INVENTION
This invention is directed generally to turbine airfoil vanes, and more particularly to hollow turbine airfoil vanes having an impingement insert for passing fluids, such as air, to cool the airfoils.
BACKGROUND OF THE INVENTION
Typically, gas turbine engines 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 vane and blade assemblies to these high temperatures. As a result, turbine vanes and blades must be made of materials capable of
withstanding such high temperatures. In addition, turbine vanes and blades often contain cooling systems for prolonging the life of the vanes and blades and reducing the likelihood of failure as a result of excessive temperatures.
Typically, turbine vanes are formed from an elongated portion forming a vane having one end configured to be coupled to a vane carrier and an opposite end configured to be movably coupled to an inner endwall. The vane is ordinarily composed of a leading edge, a trailing edge, a suction side, and a pressure side. The inner aspects of most turbine vanes typically contain an intricate maze of cooling circuits forming a cooling system. The cooling circuits in the vanes receive cooling fluid, e.g., air from the compressor of the turbine engine, and pass the fluid through the ends of the vane adapted to be coupled to the vane carrier. The cooling circuits often include multiple flow paths that are designed to maintain all aspects of the turbine vane at a relatively uniform temperature. At least some of the fluid passing through these cooling circuits is exhausted through orifices in the leading edge, trailing edge, suction side, and pressure side of the vane.
The cooling system, as shown in FIG. 3, may include an impingement plate 3 with a plurality of impingement holes 4 for directing cooling fluids to impinge on the outer wall 6 forming a turbine airfoil. The impingement plate 3 may be offset from the outer wall 6 a conventional distance. The impingement plate 3 may be generally flat and reside in a single plane. In this configuration, the cross flow of cooling fluids often disrupts the impingement jets directed towards the outer wall, thereby negatively impacting the cooling function of the impingement jets. While advances have been made in the cooling systems in turbine vanes, a need still exists for a turbine vane having increased cooling efficiency for dissipating heat and passing a sufficient amount of cooling fluid through the vane.
SUMMARY OF THE INVENTION
In accordance with a first aspect of the invention, a turbine vane is provided comprising a generally elongated hollow airfoil and a cooling system. The airfoil comprises an outer wall including a leading edge, a trailing edge, a pressure side, a suction side, an outer endwall at a first end, and an inner endwall at a second end opposite the first end. The cooling system is positioned within the airfoil and comprises a cooling chamber and an impingement insert positioned in the cooling chamber. The impingement insert and an inner surface of the airfoil outer wall define a cooling channel therebetween. The impingement insert includes a plurality of impingement nozzles extending toward the inner surface of the outer wall and a plurality of impingement orifices. At least one of the impingement orifices is arranged in a non-aligned pattern with respect to at least one adjacent impingement orifice such that cooling fluid passing out of the at least one impingement orifice does not directly flow into a centerline of a cooling fluid flowpath of cooling fluid passing out of the at least one adjacent
impingement orifice.
Each impingement orifice may be arranged in a non-aligned pattern with respect to at least one adjacent impingement orifice such that cooling fluid passing out of each impingement orifice does not directly flow into a centerline of a cooling fluid flowpath of cooling fluid passing out of the at least one adjacent impingement orifice.
The orifices may be arranged in a staggered pattern comprising alternating first and second rows that are offset from one another in a flow direction of cooling fluid through the cooling channel.
Each of the impingement orifices may be formed in an outermost aspect of a corresponding impingement nozzle, and the impingement nozzles may have a generally cylindrical cross sectional area. Each impingement nozzle may include at least one impingement orifice for directing cooling fluids orthogonally away from the impingement insert. A distance between the outermost aspect of the impingement nozzle and the inner surface of the outer wall may be less than half of a distance between an innermost aspect of the impingement insert and the inner surface of the outer wall.
Only select ones of the impingement nozzles may include a corresponding impingement orifice formed therein.
In accordance with a second aspect of the invention, a turbine vane is provided comprising a generally elongated hollow airfoil and a cooling system. The airfoil comprises an outer wall including a leading edge, a trailing edge, a pressure side, a suction side, an outer endwall at a first end, and an inner endwall at a second end opposite the first end. The cooling system is positioned within the airfoil and comprises a cooling chamber and an impingement insert positioned in the cooling chamber. The impingement insert and an inner surface of the airfoil outer wall define a cooling channel therebetween. The impingement insert includes a plurality of impingement nozzles extending toward the inner surface of the outer wall and a plurality of impingement orifices. The impingement orifices are arranged in a staggered pattern comprising alternating first and second rows that are offset from one another in a flow direction of cooling fluid through the cooling channel such that cooling fluid passing out of each respective impingement orifice does not directly flow into a centerline of a cooling fluid flowpath of cooling fluid passing out of the adjacent upstream impingement orifice.
In accordance with a third aspect of the invention a turbine vane is provided comprising a generally elongated hollow airfoil and a cooling system. The airfoil comprises an outer wall including a leading edge, a trailing edge, a pressure side, a suction side, an outer endwall at a first end, and an inner endwall at a second end opposite the first end. The cooling system is positioned within the airfoil and comprises a cooling chamber and an impingement insert positioned in the cooling chamber. The impingement insert and an inner surface of the airfoil outer wall define a cooling channel therebetween. The impingement insert includes a plurality of impingement nozzles extending toward the inner surface of the outer wall and a plurality of impingement orifices. The impingement orifices are formed in corresponding impingement nozzles and are arranged in a non-aligned pattern with respect to at least one adjacent impingement orifice such that cooling fluid passing out of the impingement orifices does not directly flow into a centerline of a cooling fluid flowpath of cooling fluid passing out of the at least one adjacent impingement orifice. A distance between an outermost aspect of at least one impingement nozzle including an impingement orifice and the inner surface of the outer wall is less than half of a distance between an innermost aspect of the impingement insert and the inner surface of the outer wall.
BRIEF DESCRIPTION OF THE DRAWINGS
While the specification concludes with claims particularly pointing out and distinctly claiming the present invention, it is believed that the present invention will be better understood from the following description in conjunction with the accompanying Drawing Figures, in which like reference numerals identify like elements, and wherein: FIG. 1 is a perspective view of a turbine vane having features according to the instant invention.
FIG. 2 is a cross-sectional view of the turbine vane shown in FIG. 1 taken along line 2-2.
FIG. 3 is a cross-sectional, detailed view taken of an outer wall of a conventional turbine airfoil with an impingement insert.
FIG. 4 is a cross-sectional, detailed view taken at detail line 4-4 in FIG. 2 displaying an impingement insert with a plurality of impingement nozzles.
FIG. 5 is a partial view of the inner surface of the outer wall taken along line 5-5 in FIG. 4 showing the impingement jets striking the outer wall and cross flow flowing therebetween.
FIGS. 6 and 7 are views similar to those of FIGS. 4 and 5 illustrating an impingement insert in accordance with another embodiment of the instant invention.
DETAILED DESCRIPTION OF THE INVENTION
In the following detailed description of the preferred embodiments, reference is made to the accompanying drawings that form a part hereof, and in which is shown by way of illustration, and not by way of limitation, specific preferred embodiments in which the invention may be practiced. It is to be understood that other embodiments may be utilized and that changes may be made without departing from the spirit and scope of the present invention.
As shown in FIGS. 1 -5, this invention is directed to a turbine airfoil vane 10 usable in a turbine engine. The turbine vane 10 may include one or more cooling systems 12 with an impingement plate 14, also referred to herein as an impingement insert 14, having one or more impingement nozzles 16. The turbine vane impingement nozzles 16 may extend towards an outer wall 20 forming the turbine vane 10 and may reduce the mixing of cooling fluids with impingement jets 22. Instead, the nozzles 16 may terminate within close proximity of the outer wall 20, thereby reducing the effect of cooling fluid cross flow 62. The cooling system 12 may be configured to cool internal and external aspects of the turbine vane 10 usable in a turbine engine. In at least one embodiment, the turbine airfoil cooling system 12 may be configured to be included within a stationary turbine vane 10, as shown in FIGS. 1 -4. The cooling system 12 may include one or more cooling chambers 26. For instance, the cooling chambers 26 may include one or more midcord cooling chambers 28 positioned in the outer wall 20.
As shown in FIGS. 1 -2, the turbine vane 10 may be formed from a generally elongated hollow airfoil 30 having an outer surface 32 adapted for use, for example, in an axial flow turbine engine. Outer surface 32 may have a generally concave shaped portion forming the pressure side 34 and a generally convex shaped portion forming the suction side 36. The turbine vane 10 may also include an outer endwall 38 at a first end 40 adapted to be coupled to a hook attachment and may include an inner endwall 42 at a second end 44. The airfoil 30 may also include a leading edge 46 and a trailing edge 48 opposite the leading edge 46.
As shown in FIG. 2, the turbine vane 10 may include an impingement insert 14 positioned in internal aspects of a central cooling chamber 26 of the cooling system 12. The impingement insert 14 may include a plurality of impingement nozzles 16, as shown in FIG. 4, extending from the impingement insert 14. In at least one embodiment, the nozzles 16 may extend toward an inner surface 52 of the outer wall 20 from the impingement insert 14. One or more of the nozzles 16 may include one or more impingement orifices 54. Each nozzle 16 may include at least one impingement orifice 54 positioned at an outermost aspect 56 of the nozzle 16 for directing cooling fluids orthogonally away from the impingement insert 14.
In at least one embodiment, one or more impingement nozzles 16 may be generally cylindrical. As such, a plurality of impingement nozzles 16 may be generally cylindrical. In other embodiments, one or more impingement nozzles 16 may have a cross-sectional area formed as a cylinder, a rectangle, a triangle, a semicircle, and other appropriate shapes. The impingement nozzles 16 may also be configured with a conical shape such that a cross-sectional area at a base 58 is greater than a cross- sectional area at the outermost aspect 56. One or a plurality of impingement nozzles 16 may be configured have a generally conical shape and may include one or more impingement orifices 54.
The impingement nozzles 16 may be aligned into rows, as shown in FIG. 5 through depiction of the impingement jets 60. The rows may extend in a generally spanwise direction, in a generally chordwise direction or other appropriate direction. Adjacent rows may be offset from each other as will be discussed in greater detail below.
The outermost aspect 56 of the impingement nozzle 16 and the impingement orifices 54 may be located a distance 64 that is less than a conventional distance 8 between a conventional impingement plate 3 with holes 4 and an outer wall 6 of a conventional vane, see FIGS. 3 and 4. In addition, a distance between an innermost aspect 68 of the impingement insert 14 and the inner surface 52 of the outer wall 20 is greater than a conventional distance 8 between the conventional impingement plate 3 with holes 4 and the outer wall 6 of the conventional vane, and is greater than twice the distance 64 between the outermost aspect 56 of the impingement nozzle 16 and the inner surface 52 of the outer wall 20, i.e., the distance 64 between the outermost aspect 56 of the impingement nozzle 16 and the inner surface 52 of the outer wall 20 is less than half of the distance between the innermost aspect 68 of the impingement insert 14 and the inner surface 52 of the outer wall 20. In such a configuration, the cross- sectional areas between the impingement nozzles 16 and the outer wall 20 is less than the distance 64 between the impingement insert 14 and the outer wall 20. Thus, the impingement nozzles 16 may be placed in closer position relative to outer wall 20 without changing overall volume of cooling fluid flow through a cooling channel 70 formed between the outer wall 20 and the impingement insert 14.
As shown in FIGS. 4 and 5, during use, cooling fluids may flow from a cooling fluid supply source (not shown) into the cooling system 12. The cooling fluids may be passed into the cooling channel 70 formed between the impingement insert 14 and the outer wall 20 through the impingement nozzles 16 and the impingement orifices 54 positioned at the outermost aspect 56 of the nozzles 16. The size and configuration of the jet 60 of cooling fluids flowing from the nozzle 16 is controlled by the shape and size of the nozzle 16. In at least one embodiment in which the nozzles 16 are generally circular, the impingement jets 60 may be generally circular when the cooling fluids strike the inner surface 52 of the outer wall 20, as shown in FIG. 5. After the cooling fluids impinge upon the inner surface 52 of the outer wall 20, the cooling fluids form a cross flow 62 flowing generally along the outer wall 20. Because the nozzles 16 extend to within close proximity of the outer wall 20, the impingement jets 60 have sufficient velocity such that the cross flow 62 does not disrupt the impingement jets 60. The cooling fluids flowing from the impingement nozzles 16 reduce the temperature of the outer wall 20.
Referring now to FIGS. 6 and 7, in accordance with another aspect of the present invention, where like structure to that of FIGS. 1 -5 includes the same reference number increased by 100, an impingement insert 1 14 and an inner surface 152 of an airfoil outer wall 120 define a cooling channel 170 therebetween. The general structure of the airfoil outer wall 120 according to this aspect of the invention may be similar to that of the airfoil outer wall 20 as described above with reference to FIGS. 1 -5.
The impingement insert 1 14 according to this aspect of the invention is part of a cooling system 1 12 and includes a plurality of impingement nozzles 1 16 extending toward the inner surface 152 of the airfoil outer wall 120. Similar to the impingement nozzles 16 discussed above, the impingement nozzles 1 16 may have a generally cylindrical cross sectional area, as shown in FIG. 6.
Select ones or all of the impingement nozzles 1 16 according to this aspect of the invention include at least one impingement orifice 154 located at an outermost aspect 156 of the impingement nozzle 1 16 for directing cooling fluid CF orthogonally away from the impingement insert 1 14, see FIG. 6. As shown in FIG. 6, a distance Di between the outermost aspect 156 of the impingement nozzles 1 16 and the inner surface 152 of the airfoil outer wall 120 may be less than half of a distance D2 between an innermost aspect 1 14A of the impingement insert 1 14 and the inner surface 152 of the outer wall 120. Hence, cooling fluid CF is discharged from the impingement nozzles 1 1 6 generally close to the inner surface 1 52 of the outer wall 120 to maximize impingement cooling of the outer wall 120 provided by the cooling fluid CF and to reduce disruption of the post impingement cooling fluid CF flowing normal to the impinging jets 1 60 within the cooling channel 1 70, as will be discussed further below.
Referring now to FIG. 7, at least one of the impingement orifices 1 54 (and all of the impingement orifices 1 54 shown in Fig. 7) is arranged in a non-aligned pattern with respect to at least one adjacent impingement orifice 154. For example, in the exemplary configuration shown in FIG. 7, the impingement orifices 1 54 are arranged in a staggered pattern comprising alternating first and second rows Ri , R2 that are offset from one another in a flow direction FD of the post impingement cooling fluid CF flowing in parallel to the outer wall 120 through the cooling channel 1 70. Hence, cooling fluid CF passing out of the at least one impingement orifice 1 54 does not directly flow into a centerline CCF of a cooling fluid flowpath FCF of cooling fluid CF passing out of the at least one adjacent impingement orifice 1 54.
As noted above, only select ones of the impingement nozzles 1 1 6 according to this aspect of the invention may include a corresponding impingement orifice 1 54 formed therein. Ones of the impingement nozzles 1 1 6 that do not include an impingement orifice could be provided to effect a more turbulent flow of cooling fluid CF through the cooling channel 1 70 to increase cooling provided to the outer wall 120 by the cooling fluid CF.
As shown in FIGS. 6 and 7, during use, cooling fluid CF may flow from a cooling fluid supply source (not shown) into the cooling system 1 12. The cooling fluid CF may be passed into the cooling channel 1 70 of the cooling system 1 12 between the impingement insert 1 14 and the outer wall 120 through the impingement nozzles 1 1 6 and the impingement orifices 1 54 positioned at the outermost aspect 1 56 of the nozzles 1 1 6. The size and configuration of the jets 1 60 of the cooling fluid CF discharged from the nozzles 1 1 6 is controlled by the shape and size of the nozzles 1 1 6. In at least one embodiment in which the nozzles 1 1 6 are generally circular, the impingement jets 1 60 may be generally circular when the cooling fluid CF strikes the inner surface 152 of the outer wall 120, as shown in FIG. 7. After the cooling fluid CF impinges upon the inner surface 152 of the outer wall 120, the cooling fluid CF flows through the cooling channel 170 along the cooling fluid flowpath FCF generally along the outer wall 120. Because the nozzles 1 16 extend to within close proximity of the outer wall 120, the impingement jets 160 have sufficient velocity such that the cooling fluid CF flowing along their respective cooling fluid flowpaths FCF does not disrupt the impingement jets 160. Additionally, less disruption of the jets 160 and of the cooling fluid flowpaths FCF is effected by the nozzles 1 16 of the present aspect of the invention, since the cooling fluid CF passing out of the impingement orifices 154 does not directly flow into the centerlines CCF of the cooling fluid flowpaths FCF of the cooling fluid CF passing out of the adjacent impingement orifices 154. The cooling fluid CF flowing from the impingement nozzles 1 16 reduces the temperature of the outer wall 120.
While particular embodiments of the present invention have been illustrated and described, it would be obvious to those skilled in the art that various other changes and modifications can be made without departing from the spirit and scope of the invention. It is therefore intended to cover in the appended claims all such changes and
modifications that are within the scope of this invention.

Claims

CLAIMS What is claimed is:
1. A turbine vane, comprising:
a generally elongated hollow airfoil comprising an outer wall, the outer wall including a leading edge, a trailing edge, a pressure side, a suction side, an outer endwall at a first end, and an inner endwall at a second end opposite the first end; and a cooling system positioned within the airfoil, the cooling system comprising a cooling chamber and an impingement insert positioned in the cooling chamber, the impingement insert and an inner surface of the airfoil outer wall defining a cooling channel therebetween, the impingement insert including:
a plurality of impingement nozzles extending toward the inner surface of the outer wall; and
a plurality of impingement orifices, at least one of the impingement orifices being arranged in a non-aligned pattern with respect to at least one adjacent impingement orifice such that cooling fluid passing out of the at least one impingement orifice does not directly flow into a centerline of a cooling fluid flowpath of cooling fluid passing out of the at least one adjacent impingement orifice.
2. The turbine vane of claim 1 , wherein each impingement orifice is arranged in a non- aligned pattern with respect to at least one adjacent impingement orifice such that cooling fluid passing out of each impingement orifice does not directly flow into a centerline of a cooling fluid flowpath of cooling fluid passing out of the at least one adjacent impingement orifice.
3. The turbine vane of claim 2, wherein the orifices are arranged in a staggered pattern comprising alternating first and second rows that are offset from one another in a flow direction of cooling fluid through the cooling channel.
4. The turbine vane of claim 1 , wherein each of the impingement orifices is formed in an outermost aspect of a corresponding impingement nozzle.
5. The turbine vane of claim 4, wherein the impingement nozzles have a generally cylindrical cross sectional area.
6. The turbine vane of claim 4, wherein each impingement nozzle includes at least one impingement orifice for directing cooling fluids orthogonally away from the impingement insert.
7. The turbine vane of claim 4, wherein a distance between the outermost aspect of the impingement nozzle and the inner surface of the outer wall is less than half of a distance between an innermost aspect of the impingement insert and the inner surface of the outer wall.
8. The turbine vane of claim 1 , wherein only select ones of the impingement nozzles include a corresponding impingement orifice formed therein.
9. A turbine vane, comprising:
a generally elongated hollow airfoil comprising an outer wall, the outer wall including a leading edge, a trailing edge, a pressure side, a suction side, an outer endwall at a first end, and an inner endwall at a second end opposite the first end; and a cooling system positioned within the airfoil, the cooling system comprising a cooling chamber and an impingement insert positioned in the cooling chamber, the impingement insert and an inner surface of the airfoil outer wall defining a cooling channel therebetween, the impingement insert including:
a plurality of impingement nozzles extending toward the inner surface of the outer wall; and a plurality of impingement orifices arranged in a staggered pattern comprising alternating first and second rows that are offset from one another in a flow direction of cooling fluid through the cooling channel such that cooling fluid passing out of each respective impingement orifice does not directly flow into a centerline of a cooling fluid flowpath of cooling fluid passing out of the adjacent upstream impingement orifice.
10. The turbine vane of claim 9, wherein each of the impingement orifices is formed in an outermost aspect of a corresponding impingement nozzle.
1 1. The turbine vane of claim 10, wherein only select ones of the impingement nozzles include a corresponding impingement orifice formed therein.
12. The turbine vane of claim 10, wherein the impingement nozzles have a generally cylindrical cross sectional area.
13. The turbine vane of claim 10, wherein each impingement nozzle includes at least one impingement orifice for directing cooling fluids orthogonally away from the impingement insert.
14. The turbine vane of claim 10, wherein a distance between the outermost aspect of the impingement nozzle and the inner surface of the outer wall is less than half of a distance between an innermost aspect of the impingement insert and the inner surface of the outer wall.
15. A turbine vane, comprising:
a generally elongated hollow airfoil comprising an outer wall, the outer wall including a leading edge, a trailing edge, a pressure side, a suction side, an outer endwall at a first end, and an inner endwall at a second end opposite the first end; and a cooling system positioned within the airfoil, the cooling system comprising a cooling chamber and an impingement insert positioned in the cooling chamber, the impingement insert and an inner surface of the airfoil outer wall defining a cooling channel therebetween, the impingement insert including:
a plurality of impingement nozzles extending toward the inner surface of the outer wall; and
a plurality of impingement orifices formed in corresponding impingement nozzles and being arranged in a non-aligned pattern with respect to at least one adjacent impingement orifice such that cooling fluid passing out of the
impingement orifices does not directly flow into a centerline of a cooling fluid flowpath of cooling fluid passing out of the at least one adjacent impingement orifice, wherein a distance between an outermost aspect of at least one impingement nozzle including an impingement orifice and the inner surface of the outer wall is less than half of a distance between an innermost aspect of the impingement insert and the inner surface of the outer wall.
16. The turbine vane of claim 15, wherein the orifices are arranged in a staggered pattern comprising alternating first and second rows that are offset from one another in a flow direction of cooling fluid through the cooling channel.
17. The turbine vane of claim 15, wherein each impingement nozzle includes at least one impingement orifice for directing cooling fluids orthogonally away from the impingement insert.
18. The turbine vane of claim 17, wherein the impingement nozzles have a generally cylindrical cross sectional area.
19. The turbine vane of claim 17, wherein each of the impingement orifices is formed in an outermost aspect of a corresponding impingement nozzle.
PCT/US2014/070702 2013-12-19 2014-12-17 Turbine airfoil vane with an impingement insert having a plurality of impingement nozzles WO2015095253A1 (en)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US14/133,773 US9347324B2 (en) 2010-09-20 2013-12-19 Turbine airfoil vane with an impingement insert having a plurality of impingement nozzles
US14/133,773 2013-12-19

Publications (1)

Publication Number Publication Date
WO2015095253A1 true WO2015095253A1 (en) 2015-06-25

Family

ID=52278845

Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/US2014/070702 WO2015095253A1 (en) 2013-12-19 2014-12-17 Turbine airfoil vane with an impingement insert having a plurality of impingement nozzles

Country Status (1)

Country Link
WO (1) WO2015095253A1 (en)

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3290639A1 (en) * 2016-09-06 2018-03-07 United Technologies Corporation Impingement cooling with increased cross-flow area
WO2019180382A1 (en) * 2018-03-23 2019-09-26 Safran Helicopter Engines Turbine stator vane cooled by air-jet impacts

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5352091A (en) * 1994-01-05 1994-10-04 United Technologies Corporation Gas turbine airfoil
US5586866A (en) * 1994-08-26 1996-12-24 Abb Management Ag Baffle-cooled wall part
US6439846B1 (en) * 1997-07-03 2002-08-27 Alstom Turbine blade wall section cooled by an impact flow
US20120070302A1 (en) * 2010-09-20 2012-03-22 Ching-Pang Lee Turbine airfoil vane with an impingement insert having a plurality of impingement nozzles

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5352091A (en) * 1994-01-05 1994-10-04 United Technologies Corporation Gas turbine airfoil
US5586866A (en) * 1994-08-26 1996-12-24 Abb Management Ag Baffle-cooled wall part
US6439846B1 (en) * 1997-07-03 2002-08-27 Alstom Turbine blade wall section cooled by an impact flow
US20120070302A1 (en) * 2010-09-20 2012-03-22 Ching-Pang Lee Turbine airfoil vane with an impingement insert having a plurality of impingement nozzles

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3290639A1 (en) * 2016-09-06 2018-03-07 United Technologies Corporation Impingement cooling with increased cross-flow area
WO2019180382A1 (en) * 2018-03-23 2019-09-26 Safran Helicopter Engines Turbine stator vane cooled by air-jet impacts
FR3079262A1 (en) * 2018-03-23 2019-09-27 Safran Helicopter Engines FIXED WATER TURBINE COOLING BY AIR JET IMPACTS
US11333025B2 (en) 2018-03-23 2022-05-17 Safran Helicopter Engines Turbine stator blade cooled by air-jet impacts

Similar Documents

Publication Publication Date Title
US9347324B2 (en) Turbine airfoil vane with an impingement insert having a plurality of impingement nozzles
US20120070302A1 (en) Turbine airfoil vane with an impingement insert having a plurality of impingement nozzles
JP6526166B2 (en) Vane cooling structure
EP2131108B1 (en) Counter-vortex film cooling hole design
EP2604800B1 (en) Nozzle vane for a gas turbine engine
US7780413B2 (en) Turbine airfoil with near wall inflow chambers
US20100221121A1 (en) Turbine airfoil cooling system with near wall pin fin cooling chambers
EP3167159B1 (en) Impingement jet strike channel system within internal cooling systems
EP3271554B1 (en) Internal cooling system with converging-diverging exit slots in trailing edge cooling channel for an airfoil in a turbine engine
US8985949B2 (en) Cooling system including wavy cooling chamber in a trailing edge portion of an airfoil assembly
JP6407276B2 (en) Gas turbine engine component including trailing edge cooling using impingement angled to a surface reinforced by a cast chevron array
US20150198050A1 (en) Internal cooling system with corrugated insert forming nearwall cooling channels for airfoil usable in a gas turbine engine
US9863256B2 (en) Internal cooling system with insert forming nearwall cooling channels in an aft cooling cavity of an airfoil usable in a gas turbine engine
US7300242B2 (en) Turbine airfoil with integral cooling system
JP6239163B2 (en) Turbine blade cooling system with leading edge impingement cooling system and adjacent wall impingement system
US20180045059A1 (en) Internal cooling system with insert forming nearwall cooling channels in an aft cooling cavity of a gas turbine airfoil including heat dissipating ribs
US8920122B2 (en) Turbine airfoil with an internal cooling system having vortex forming turbulators
JP2015105656A (en) Turbine blade with near wall microcircuit edge cooling
JP6203400B2 (en) Turbine blade with a laterally extending snubber having an internal cooling system
WO2015095253A1 (en) Turbine airfoil vane with an impingement insert having a plurality of impingement nozzles
WO2015195088A1 (en) Turbine airfoil cooling system with leading edge impingement cooling system
WO2015134006A1 (en) Turbine blade with film cooling leading edge showerhead
WO2016133513A1 (en) Turbine airfoil with a segmented internal wall
WO2015191037A1 (en) Turbine airfoil cooling system with leading edge diffusion film cooling holes

Legal Events

Date Code Title Description
121 Ep: the epo has been informed by wipo that ep was designated in this application

Ref document number: 14822029

Country of ref document: EP

Kind code of ref document: A1

NENP Non-entry into the national phase

Ref country code: DE

122 Ep: pct application non-entry in european phase

Ref document number: 14822029

Country of ref document: EP

Kind code of ref document: A1