WO2015191037A1 - Système de refroidissement d'une aube de turbine avec trous de refroidissement par film de diffusion de bord d'attaque - Google Patents

Système de refroidissement d'une aube de turbine avec trous de refroidissement par film de diffusion de bord d'attaque Download PDF

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Publication number
WO2015191037A1
WO2015191037A1 PCT/US2014/041619 US2014041619W WO2015191037A1 WO 2015191037 A1 WO2015191037 A1 WO 2015191037A1 US 2014041619 W US2014041619 W US 2014041619W WO 2015191037 A1 WO2015191037 A1 WO 2015191037A1
Authority
WO
WIPO (PCT)
Prior art keywords
airfoil
diffusion film
section
film cooling
leading edge
Prior art date
Application number
PCT/US2014/041619
Other languages
English (en)
Inventor
Nan Jiang
Ralph W. Matthews
Robert A. Dorris
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/US2014/041619 priority Critical patent/WO2015191037A1/fr
Publication of WO2015191037A1 publication Critical patent/WO2015191037A1/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
    • F05D2240/10—Stators
    • F05D2240/12—Fluid guiding means, e.g. vanes
    • F05D2240/121—Fluid guiding means, e.g. vanes related to the leading edge of a stator vane
    • 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
    • F05D2250/00—Geometry
    • F05D2250/10—Two-dimensional
    • F05D2250/12—Two-dimensional rectangular
    • 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
    • F05D2250/00—Geometry
    • F05D2250/20—Three-dimensional
    • F05D2250/21—Three-dimensional pyramidal
    • 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
    • F05D2250/00—Geometry
    • F05D2250/30—Arrangement of components
    • F05D2250/32—Arrangement of components according to their shape
    • F05D2250/324—Arrangement of components according to their shape divergent
    • 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
    • F05D2250/00—Geometry
    • F05D2250/70—Shape
    • F05D2250/71—Shape curved
    • 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
    • F05D2250/00—Geometry
    • F05D2250/70—Shape
    • F05D2250/71—Shape curved
    • F05D2250/711—Shape curved convex
    • 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

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 one or more diffusion film cooling holes with an exhaust outlet positioned at the stagnation line at the leading edge and configured to exhaust cooling fluid to the pressure and suction sides of the airfoil is disclosed.
  • the diffusion film cooling hole may be formed from a first section having a generally constant cross-section and a second section extending outward from the first section with a diverging cross-sectional area.
  • the exhaust outlet of the diffusion film cooling hole may include a curved side that follows the curvature of the outer surface at the leading edge of the airfoil.
  • the turbine airfoil may include a showerhead at the leading edge formed from a single row of diffusion film cooling holes that exhaust cooling fluid to the pressure and suction sides of the airfoil.
  • the turbine airfoil may be formed from a generally elongated hollow airfoil formed from an outer wall, and having a leading edge, a trailing edge, a pressure side, a suction side, and a cooling system positioned within interior aspects of the generally elongated hollow airfoil and formed from at least one cavity.
  • the cooling system may include one or more diffusion film cooling holes positioned in the outer wall forming the generally elongated airfoil and providing a cooling fluid pathway between the at least one cavity forming the cooling system and an environment outside of the airfoil.
  • the diffusion film cooling hole may include a first section extending from an inner surface of the outer wall into the outer wall and a second section extending the first section and terminating at an outer surface of the outer wall.
  • One or more surfaces defining the second section may extend outwardly from an intersection of the first and second sections towards the outer surface such that the surface is angled way from a longitudinal axis of the diffusion film cooling hole thereby increasing a size of a cross-sectional area of the second section.
  • An exhaust outlet of the diffusion film cooling hole at an intersection of the second section and the outer surface of the outer wall may include first and second sides, whereby the first side has a generally curved side.
  • the first and second sides may define opposite sides of the exhaust outlet and may be longer than third and fourth sides that define opposite sides of the exhaust outlet and extend between the first and second sides.
  • the exhaust outlet may extend chordwise from a stagnation line on the leading edge into the pressure side and from the stagnation line into the suction side.
  • the first side may have a generally curved side that is curved about a curvature of the outer surface at the leading edge of the airfoil.
  • the curved side of the first side may curve from the third side to the fourth side.
  • the first side may diverge from the longitudinal axis between about 5 degrees and about 30 degrees. In another embodiment, the first side may diverges from the longitudinal axis less than about 10 degrees.
  • the second side may be parallel with the longitudinal axis.
  • the third side may diverge from the longitudinal axis moving away from the intersection of the first and second sections. In at least one embodiment, the third side may diverge from the longitudinal axis moving away from the intersection of the first and second sections between about 5 degrees and about 30 degrees.
  • the fourth side may diverge from the longitudinal axis moving away from the intersection of the first and second sections. In at least one embodiment, the fourth side may diverge from the longitudinal axis moving away from the intersection of the first and second sections between about 5 degrees and about 30 degrees.
  • the diffusion film cooling hole may include a plurality of one diffusion film cooling holes aligned into a row extending spanwise along the stagnation line at the leading edge of the airfoil.
  • cooling fluids such as, but not limited to, air
  • the cooling fluid may enter the cavity and flow into the first section of a diffusion film cooling hole.
  • the cooling fluid flows through the first section with a generally consistent cross-sectional area.
  • the cooling fluid passed into the second section with a diverging cross-sectional area.
  • the cooling fluids are exhausted from the exhaust outlet at the stagnation line and emitted at the pressure side and the suction side from the same exhaust outlet because the exhaust outlet is positioned at the stagnation line and extends from the pressure side to the suction side.
  • An advantage of the diffusion film cooling hole is that the diffusion film cooling hole exhausts cooling fluid to form film cooling on the pressure side and suction side of the airfoil, thereby accomplishing in a single row diffusion film cooling holes what took multiple rows of conventional film cooling holes.
  • diffusion film cooling hole Another advantage of the diffusion film cooling hole is that manufacturing the diffusion film cooling holes aligned into a row along the stagnation line at the leading edge is cheaper than forming multiple rows of conventional film cooling holes at a leading edge of an airfoil to form a showerhead.
  • diffusion film cooling hole provides better film cooling coverage than conventional straight film cooling holes.
  • diffusion film cooling hole is that because the diffusion film cooling hole is positioned at the stagnation line, a greater amount of cooling fluid will be emitted at the stagnation line compared to conventional film cooling holes, thereby enabling a pitch of the diffusion film cooling hole to be reduced for better film coverage in comparison to convention film holes.
  • Figure 1 is a perspective view of a turbine airfoil with an internal cooling system including diffusion cooling holes at the stagnation line on the leading edge.
  • FIG 2 is another perspective view of the turbine airfoil of Figure 1 with an internal cooling system including diffusion cooling holes at the stagnation line on the leading edge.
  • FIG 3 is yet another perspective view of the turbine airfoil of Figure 1 with an internal cooling system including diffusion cooling holes at the stagnation line on the leading edge.
  • Figure 4 is a rear view of a diffusion film cooling hole.
  • Figure 5 is a side perspective view of the diffusion film cooling hole of Figure
  • Figure 6 is a front view of the diffusion film cooling hole of Figure 4.
  • a turbine airfoil 10 usable in a turbine engine and having an internal cooling system 14 with one or more diffusion film cooling holes 16 with an exhaust outlet 18 positioned at the stagnation line 20 at the leading edge 22 and configured to exhaust cooling fluid to the pressure and suction sides 24, 26 of the airfoil 10 is disclosed.
  • the diffusion film cooling hole 16 may be formed from a first section 28 having a generally constant cross-section and a second section 30 extending outward from the first section 28 with a diverging cross-sectional area.
  • the exhaust outlet 18 of the diffusion film cooling hole 16 may include a curved side 32 that follows the curvature of the outer surface 34 at the leading edge 22 of the airfoil 10.
  • the turbine airfoil 10 may include a
  • the turbine airfoil 10 may be formed from a generally elongated hollow airfoil 40 formed from an outer wall 42 and having a leading edge 22, a trailing edge 44, a pressure side 24, a suction side 26, and a cooling system 14 positioned within interior aspects of the generally elongated hollow airfoil 40 and formed from one cavities 48.
  • the turbine airfoil 10 may be a turbine blade, which rotates with the rotor assembly, or a turbine vane, which remained fixed during steady state operation.
  • the outer wall 42 forming the generally elongated airfoil 40 may have one or more diffusion film cooling holes 16 positioned in the outer wall 42 and providing a cooling fluid pathway between the cavity 48 forming a portion of the internal cooling system 14 and an environment outside of the airfoil 40.
  • the diffusion film cooling hole 16 may include a first section 30 extending from an inner surface 50 of the outer wall 42 into the outer wall 42 and a second section 30 extending the first section 28 and terminating at an outer surface 34 of the outer wall 42.
  • a surface 54 defining the second section 30 may extend outwardly from an intersection 56 of the first and second sections 28, 30 towards the outer surface 34 such that the surface 54 is angled way from a longitudinal axis 58 of the diffusion film cooling hole 16 thereby increasing a size of a cross-sectional area of the second section 30.
  • the diffusion film cooling hole 16 may include an exhaust outlet 18 at an intersection 62 of the second section 30 and the outer surface 34 of the outer wall 42.
  • the exhaust outlet 18 may include first and second sides 64, 66, whereby the first side 64 has a generally curved side.
  • the first and second sides 64, 66 may define opposite sides of the exhaust outlet 18 and may be longer than third and fourth sides 68, 70 that define opposite sides of the exhaust outlet 18 and extend between the first and second sides 64, 66.
  • the exhaust outlet 18 may extend chordwise from the stagnation line 20 on the leading edge 22 into the pressure side 24 and from the stagnation line 20 into the suction side 26. In at least one
  • the exhaust outlet 18 may be between about 3 times and about 15 times larger than an inlet 74 of the first section 28 or the inlet 76 of the second section 30.
  • the first side 64 may have a generally curved side that is curved about a curvature of the outer surface 34 at the leading edge 22 of the airfoil 40. In at least one embodiment, as shown in Figures 4-6, the curved side of the first side 64 may extend from the third side 68 to the fourth side 70. The first side 64 may diverge from the longitudinal axis 58 between about 5 degrees and about 30 degrees. In another embodiment, the first side 64 may diverge from the longitudinal axis 58 less than about 10 degrees.
  • the second side 66 may be parallel with the longitudinal axis 58. The second side 66 may be generally linear.
  • the third side 68 may diverge from the longitudinal axis 58 moving away from the intersection 56 of the first and second sections 28, 30. In one embodiment, the third side 68 may diverge from the longitudinal axis 58 moving away from the intersection 56 of the first and second sections 28, 30 between about 5 degrees and about 30 degrees.
  • the third side is generally linear.
  • the fourth side 70 may diverge from the longitudinal axis 58 moving away from the intersection 56 of the first and second sections 28, 30.
  • the fourth side 70 may diverge from the longitudinal axis 58 moving away from the intersection 56 of the first and second sections 28, 30 between about 5 degrees and about 30 degrees.
  • the fourth side 68 may be generally linear.
  • the turbine airfoil 10 may include a plurality of diffusion film cooling holes 16 at the leading edge 22 of the airfoil 40.
  • the plurality of diffusion film cooling holes 16 may be aligned into a spanwise extending row 38 at the leading edge 22 of the airfoil 40.
  • the spanwise extending row 38 plurality of diffusion film cooling holes 16 may extend along the stagnation line 20 at the leading edge 22 of the airfoil 40.
  • cooling fluids such as, but not limited to, air
  • the cooling fluid may enter the cavity 48 and flow into the first section 28 of a diffusion film cooling hole 16.
  • the cooling fluid flows through the first section 28 with a generally consistent cross-sectional area.
  • the cooling fluids are exhausted from the exhaust outlet 18 at the stagnation line 20 and emitted at the pressure side 24 and the suction side 26 from the same exhaust outlet 18 because the exhaust outlet is positioned at the stagnation line and extends from the pressure side to the suction side.

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

Abstract

L'invention concerne une aube de turbine (10) utilisable dans un moteur à turbine et comportant un système interne de refroidissement (14) comprenant un ou plusieurs trous de refroidissement par film de diffusion (16) avec une sortie d'évacuation (18) positionnée au niveau de la ligne de stagnation (20) au niveau du bord d'attaque (22), et configurée pour évacuer le fluide de refroidissement vers les côtés pression et aspiration (24, 26) de l'aube (10). Le trou de refroidissement par film de diffusion (16) peut se composer d'une première section (28) ayant une section transversale généralement constante, et d'une seconde section (30) s'étendant vers l'extérieur depuis la première section (28) et possédant une zone de section transversale divergente. La sortie d'évacuation (18) du trou de refroidissement par film de diffusion (16) peut comprendre un côté courbe qui suit la courbure de la surface externe (34) au niveau du bord d'attaque (22). Dans au moins un mode de réalisation, l'aube de turbine (10) peut comprendre une groupe de trous (36) au niveau du bord d'attaque (22), composé d'une seule rangée (38) de trous de refroidissement par film de diffusion (16) qui évacuent le fluide de refroidissement vers les côtés pression et aspiration (24, 26) de l'aube (10).
PCT/US2014/041619 2014-06-10 2014-06-10 Système de refroidissement d'une aube de turbine avec trous de refroidissement par film de diffusion de bord d'attaque WO2015191037A1 (fr)

Priority Applications (1)

Application Number Priority Date Filing Date Title
PCT/US2014/041619 WO2015191037A1 (fr) 2014-06-10 2014-06-10 Système de refroidissement d'une aube de turbine avec trous de refroidissement par film de diffusion de bord d'attaque

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
PCT/US2014/041619 WO2015191037A1 (fr) 2014-06-10 2014-06-10 Système de refroidissement d'une aube de turbine avec trous de refroidissement par film de diffusion de bord d'attaque

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Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US10577942B2 (en) 2016-11-17 2020-03-03 General Electric Company Double impingement slot cap assembly

Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1609949A1 (fr) * 2004-06-23 2005-12-28 General Electric Company Paroi refroidie par film d'air avec trous de refroidissement en forme de chevron
EP1691033A1 (fr) * 2005-01-28 2006-08-16 The General Electric Company Trou rectangulaire en forme de diffuseur pour une aube de turbine refroidie par couche d'air
WO2007006619A1 (fr) * 2005-07-12 2007-01-18 Siemens Aktiengesellschaft Composant a refroidissement par film, en particulier une pale de turbine et procede de fabrication d'une pale de turbine
US8066484B1 (en) * 2007-11-19 2011-11-29 Florida Turbine Technologies, Inc. Film cooling hole for a turbine airfoil
US8245519B1 (en) * 2008-11-25 2012-08-21 Florida Turbine Technologies, Inc. Laser shaped film cooling hole

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1609949A1 (fr) * 2004-06-23 2005-12-28 General Electric Company Paroi refroidie par film d'air avec trous de refroidissement en forme de chevron
EP1691033A1 (fr) * 2005-01-28 2006-08-16 The General Electric Company Trou rectangulaire en forme de diffuseur pour une aube de turbine refroidie par couche d'air
WO2007006619A1 (fr) * 2005-07-12 2007-01-18 Siemens Aktiengesellschaft Composant a refroidissement par film, en particulier une pale de turbine et procede de fabrication d'une pale de turbine
US8066484B1 (en) * 2007-11-19 2011-11-29 Florida Turbine Technologies, Inc. Film cooling hole for a turbine airfoil
US8245519B1 (en) * 2008-11-25 2012-08-21 Florida Turbine Technologies, Inc. Laser shaped film cooling hole

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US10577942B2 (en) 2016-11-17 2020-03-03 General Electric Company Double impingement slot cap assembly

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