WO2014106598A1 - Aube de turbomachine - Google Patents

Aube de turbomachine Download PDF

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Publication number
WO2014106598A1
WO2014106598A1 PCT/EP2013/077812 EP2013077812W WO2014106598A1 WO 2014106598 A1 WO2014106598 A1 WO 2014106598A1 EP 2013077812 W EP2013077812 W EP 2013077812W WO 2014106598 A1 WO2014106598 A1 WO 2014106598A1
Authority
WO
WIPO (PCT)
Prior art keywords
cooling
blade
airfoil
inlet
cooling fluid
Prior art date
Application number
PCT/EP2013/077812
Other languages
English (en)
Inventor
Janos Szijarto
Esa Utriainen
Per Almroth
Magnus Hasselqvist
Lieke Wang
Original Assignee
Siemens Aktiengesellschaft
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Siemens Aktiengesellschaft filed Critical Siemens Aktiengesellschaft
Publication of WO2014106598A1 publication Critical patent/WO2014106598A1/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
    • F01D5/187—Convection cooling
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B22—CASTING; POWDER METALLURGY
    • B22C—FOUNDRY MOULDING
    • B22C9/00—Moulds or cores; Moulding processes
    • B22C9/02—Sand moulds or like moulds for shaped castings
    • B22C9/04—Use of lost patterns
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B22—CASTING; POWDER METALLURGY
    • B22C—FOUNDRY MOULDING
    • B22C9/00—Moulds or cores; Moulding processes
    • B22C9/10—Cores; Manufacture or installation of cores
    • 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
    • F05D2230/00—Manufacture
    • F05D2230/20—Manufacture essentially without removing material
    • F05D2230/21—Manufacture essentially without removing material by casting
    • F05D2230/211—Manufacture essentially without removing material by casting by precision casting, e.g. microfusing or investment casting

Definitions

  • Blade for a turbomachine The present invention relates to a blade for a turbomachine and more particularly to a cooling arrangement for an airfoil of a blade of a turbomachine.
  • the blade typically includes an airfoil portion and a root portion separated by a platform.
  • the airfoil portion of the blade is cooled by directing a cooling fluid to flow through radial passages formed in the airfoil portion of the blades.
  • a number of small axial passages are formed inside the blade airfoils that connect with one or more of the radi ⁇ al passages so that cooling air is directed over the surfaces of the airfoils, such as the leading and trailing edges or the suction and pressure surfaces. After the cooling air ex- its the blade it enters and mixes with the hot gas flowing through the turbine section.
  • cooling of the blade is achieved by supplying the cooling fluid from the compressor to the cooling channels in the blades.
  • the cooling channels often include multiple flow paths that are designed to maintain all aspects of the tur ⁇ bine blade at a relatively uniform temperature.
  • Several different cooling arrangements based on a combination of convective, impingement, and external film-based cooling have been proposed in the state of the art. It is an object of the present invention to provide an im ⁇ proved and efficient cooling arrangement for the blade of a turbomachine .
  • the object is achieved by providing a blade for a turbo- machine according to claim 1.
  • a blade for a turbomachine includes an airfoil portion and a root portion, the airfoil portion including a pressure side and a suction side extending between a lead ⁇ ing edge and a trailing edge of the airfoil portion, charac ⁇ terized in that the airfoil portion comprises a cooling ar ⁇ rangement, having an inlet system for introducing cooling fluid into the airfoil portion for cooling the airfoil por- tion, at least one first cooling passage for conducting cool ⁇ ing fluid in a direction from the trailing edge to the leading edge on the suction side, wherein the first cooling passage is fluidly connected to the inlet system and at least one second cooling passage for conducting cooling fluid in a direction from the leading edge to the trailing edge on the pressure side, wherein the second cooling passage is fluidly connected to the inlet system and wherein the cooling fluid flowing in the second cooling passage is directed out of the airfoil portion through an opening at the trail
  • the cooling efficiency is increased since the pressure side and the suction side sur ⁇ faces of the airfoil are respectively and independently cooled by the cooling fluid flowing in different and opposite directions .
  • the airfoil portion extends in a direction radial to an axis of rotation of a rotor of the turbomachine .
  • a separating cavity between the inlet sys ⁇ tem and the first cooling passage and the second cooling pas ⁇ sage enables directing portions of the cooling fluid to the passages respectively, thereby enabling cooling of both the suction side and the pressure side surfaces of the airfoil.
  • design of the separating cavity may also enable directing a first portion of the fluid in the first cooling passage and a second portion of the fluid in the second cool ⁇ ing passage depending on the amount of cooling required for the suction side and the pressure side.
  • the inlet system includes a first inlet and a second inlet fluidly connected to the first cool ⁇ ing passage and the second cooling passage respectively.
  • Such an arrangement enables a fixed and desired amount of cooling fluid into the airfoil based on the cooling required at the pressure side and the suction side.
  • the first inlet for the suction side cooling cavities may have a smaller cross sectional area than the se ⁇ cond inlet for the pressure side cooling cavities.
  • a plurality of impingement holes is pre ⁇ sent in the airfoil for cooling the leading edge portion of the blade. Cooling fluid flowing along the first cooling pas ⁇ sage is directed towards the leading edge, which is exposed to high temperatures, hence impingement holes provide im- pingement cooling at the leading edge portion, prior to the cooling fluid exiting the blade from an outlet at the leading edge .
  • the cooling passages include a plurality of cavities and/or cooling channels. Presence of cavities and channels enables directing the cooling fluid to all regions/ portions of the airfoil.
  • the cavities at the suction side are smaller than the cavities at the pressure side.
  • Such an ar ⁇ rangement enables higher velocity of fluid flowing in the cavities at the suction side resulting in improved cooling.
  • “Smaller” means particularly a smaller - i.e. reduced - cross sectional area of the cavities at the suction side compared to the cavities at the pressure side, e.g. when the cross sectional area is taken perpendicular to the cooling fluid flow within the cavity, e.g. at a specific airfoil height.
  • the blade includes a platform separating the airfoil portion and the root portion, and the inlet sys ⁇ tem is present at the platform. Cooling fluid from the root portion is directed into the airfoil portion via the inlet system located at the platform.
  • the cooling fluid flowing in the first cooling passage is directed out of the airfoil portion through a plurality of film cooling holes present at the leading edge.
  • the film cooling holes enable cooling of the external surface of the blade providing a thin, cool, insu ⁇ lating blanket of cooling fluid thereby lowering the tempera ⁇ ture .
  • the cooling fluid flowing in the second cooling passage is discharged out of the airfoil por ⁇ tion through an opening at the trailing edge.
  • the second cooling passage includes a cooling channel that has a first portion on the suction side extending in a direction from the trailing edge to the leading edge and a second portion downstream the first portion, the second portion extending from the leading edge to the trailing edge.
  • FIG. 1 is a schematic diagram of a blade of a turbomachine
  • FIG. 2 is a cross sectional view depicting a platform of the blade of FIG. 1 with an inlet system
  • FIG. 3 is an isometric view of the blade of FIG. 1,
  • FIG. 4 is a cross sectional view depicting another embodiment of the blade of FIG. 1,
  • FIG. 5 is a cross sectional view depicting yet another em ⁇ bodiment of the blade of FIG. 1, in accordance with the as ⁇ pects of the present technique.
  • Embodiments of the present invention described below relate to a blade component in a turbomachine.
  • the turbomachine may include a gas turbine, a steam turbine, a turbofan and the like .
  • FIG. 1 is a schematic diagram of an exemplary blade 1 of a rotor (not shown) of a turbomachine, such as a gas turbine and
  • FIG. 2 is a cross sectional view of the blade depicting a platform of the blade with an inlet system.
  • the blade 1 includes an air ⁇ foil portion 2 and a root portion 3.
  • the airfoil portion 2 projects from the root portion 3 in a radial direction as de ⁇ picted, wherein the radial direction means a direction per- pendicular to the rotation axis of the rotor.
  • the airfoil portion 2 extends radially along a longitudinal direc ⁇ tion of the blade 1.
  • the blade 1 is attached to a body of the rotor (not shown) , in such a way that the root portion 3 is attached to the body of the rotor whereas the airfoil portion 2 is located at a radially outermost position.
  • the airfoil portion 2 has an outer wall 10 including a pressure side 6, also called pressure surface, and a suction side 7, also called suction surface.
  • the pressure side 6 and the suction side 7 are joined together along an upstream leading edge 4 and a downstream trailing edge 5, wherein the leading edge 4 and the trailing edge 5 are spaced axially from each other as depicted in FIG. 1.
  • the outer wall portion on the pressure side may be referred to as the pressure-side wall 11 and the outer wall portion on the suction side may be referred to as the suction-side wall 12.
  • the respective surfaces of the walls 11, 12 facing the internal region are referred to as inner surfaces. Simi ⁇ larly, the respective surfaces of the walls 11, 12 facing the external region are referred to as outer surfaces.
  • one or more cooling holes 8 are present on the pressure side 6 and the suction side 7 of the blade as depicted in FIG. 1. The cooling holes 8 aid in film cooling of the blade 1 as will be described in more detail with reference to FIG. 2.
  • a platform 9 is formed at an upper portion of the root por- tion 3.
  • the airfoil portion 2 is connected to the platform 9 and extends in the radial direction outward from the platform 9.
  • the air- foil portion 2 of the blade 1 includes a cooling arrangement, which includes an intricate maze of internal structures such as cooling passages having cavities, channels and other structures such as ribs and pin fins for enabling enhanced cooling .
  • the cooling arrangement includes an inlet system 28 located at the platform between the root portion 3 and the airfoil portion 2 for introducing the cooling fluid 26 which may be cool air or a coolant into the airfoil portion 2, for cooling the airfoil portion 2.
  • the cooling fluid 26 which may be cool air or a coolant into the airfoil portion 2, for cooling the airfoil portion 2.
  • the inlet system 28 includes one or more inlets, such as a first inlet 29 and a second inlet 30 for directing the cool ⁇ ing fluid into the airfoil portion 2 of the blade 1.
  • the blade 1 may be cast as a single component or may alternatively be assembled from multiple components.
  • the multiple component blade may include a leading edge component, a trailing edge component and a core region component.
  • the components may be cast separately and thereafter joined together by bonding or brazing for example .
  • the blade is manufactured using precision casting technique.
  • the cooling fluid 26 is directed to various portions of the airfoil 2 through an intricate maze of passages.
  • at least one first cooling passage is fluidly connected to the inlet sys ⁇ tem 28 for conducting the cooling fluid in a direction from the trailing edge 5 to the leading edge 4 on the suction side 7.
  • At least one second cooling passage is fluidly connected to the inlet system 28 for conducting the cooling fluid in a direction from the leading edge 4 to the trailing edge 5 on the pressure side 6.
  • cooling fluid 26 is discharged from the blade at the leading edge 4 through the plurality of film cooling holes 8 present on the leading edge 4, whereas the cooling fluid 26 is discharged from the blade airfoil 2 at the trailing edge 5 through single or multiple opening 13 present at the trailing edge 5.
  • the airfoil 2 includes a plurality of cooling passages, wherein the cooling passages include a plurality of cavities and channels for conducting the cooling fluid 26 therein.
  • the cooling fluid may be present in a cooling-fluid source which may be located external to the blade 1.
  • the cooling-fluid source may be in ⁇ ternal to the blade 1 wherein the cooling fluid is stored in the root portion 3 of the blade 1.
  • the inlet system 28 such as the inlet system 28 of FIG. 2 located at the platform 9 directs the cooling fluid 26 inside the airfoil 2 through one or more cooling passages.
  • the inlet system 28 includes one or more inlets 29, 30 for supplying the cooling fluid 26 to the air ⁇ foil 2.
  • the cooling fluid 26 is directed into a first cooling passage 20, at the suction side 7.
  • the first cooling passage 20 includes a first trailing cavity 21, a core cavity 22 and a first cav- ity 24 in fluid communication through one or more cooling channels .
  • the blade 1 may have three regions, namely a lead ⁇ ing region, a trailing region and a core region between the leading region and the trailing region.
  • the cavities present at the leading region, core region and the trailing region are referred to as the leading cavity, core cavity and the trailing cavity respectively.
  • the airfoil portion 2 of the blade has a first end 15 and a second end 17 extending in a direction ra ⁇ dial to the root portion 3, wherein the second end 17 is at the platform 9, adjacent to the root portion 3 and the first end 15 is distal from the platform 9 and the root portion 3.
  • the first end 15 is also referred to as the tip of the blade 1.
  • first trailing cavity 21 is connected to the core cavity 22 and the core cavity 22 is connected to the first cavity 24 through channels lo ⁇ cated at the first end 15 and the second end 17 of the blade.
  • the cooling fluid 26 enters the first trailing cavity 21 via the inlet system 28 located at the platform 9 of the blade 1 through the root portion 3. Thereafter, the fluid 26 is di ⁇ rected to the core cavity 22 and subsequently to the first cavity 24. The fluid 26 after getting directed to the first cavity 24 impinges on a leading edge cavity 27 through a plurality of impingement holes 36. The fluid 26 is subsequently discharged from the blade through the film cooling holes 8 present at the leading edge 4 of the blade.
  • the cooling fluid 26 is directed into a second cooling passage 31 which includes a second cavity 32, a second core cavity 34 and a second trailing cavity 35.
  • the fluid 26 is directed into the second cavity 32 located on the pressure side 6, thereafter into the second core cavity 34 present at the core region of the blade through a channel present at a radially first end 15 opposite the root portion 3.
  • the cooling fluid 26 enters the second trailing cavity 35 through a second channel located at the second end 17 of the blade, and subsequently discharged through the opening 13 at the trailing edge 5.
  • the inlet system 28 includes two inlets, a first inlet 29 supplies the cooling fluid 26 to a first cooling passage 40 and a second inlet 30 is fluidly connected to a second cooling passage 50 for supplying the cooling fluid.
  • the first cooling passage 40 includes a cavity 42 located adjacent the suction side 7, a channel 46 direct- ing the fluid 26 from the cavity 42 into the leading edge cavity 44.
  • the second cooling passage 50 includes a core cavity 48, lo ⁇ cated at the core region wherein the core cavity 48 is fluid- ly connected to the second inlet 30.
  • the cooling channel 51 includes a first portion 52 and a second portion 54.
  • the first portion 52 is on the suction side 7 extending in a direction from the trailing edge 5 to the leading edge 4 and the second portion 54 which is downstream the first portion 52, the second portion 54 extending in a direction from the leading edge 4 to the trailing edge 5.
  • the second core cavity 56 distributes the cooling into a trailing edge cavity 58 from where the cooling fluid 26 is discharged through the opening 13 at the trailing edge 5 into the hot gas path.
  • FIG. 5 is a cross sectional view depicting yet another em ⁇ bodiment of the blade of FIG. 1.
  • Cooling fluid 26 is supplied into airfoil portion 2 through the inlet system 28, which in the present embodiment includes one inlet (not shown in FIG. 5) from the root portion 3, wherein the inlet is located at the platform 9 of the blade.
  • the cooling fluid 26 enters the core cavity 62 and turns at the tip 15 of the airfoil 2 to enter a second cavity 64.
  • the second cavity 64 is fluidly connected to a separating cavity 66 at the second end 17 of the blade located proximal to the platform 9.
  • the separating cavity 66 directs a first portion of the cool- ing fluid 26 into a first cooling passage 68 which includes the channel 70 located at the suction side 7 and the leading edge cavity 72 and directs a second portion of the cooling fluid 26 into the second cooling passage 69 which includes the trailing edge cavities 73, 74 and thereafter discharges the fluid 26 through the opening 13 at the trailing edge 5.
  • the cavities and/or channels at the suction side 7 are smaller than the cavities at the pressure side 6.

Abstract

La présente invention concerne une aube (1) de turbomachine. L'aube (1) comprend une partie de profil aérodynamique (2) et une partie d'amplanture (3), le profil aérodynamique (2) comprenant un côté de pression (6) et un côté d'aspiration (7) s'étendant entre un bord d'attaque (4) est un bord de fuite (5) de la partie de profil aérodynamique (2), caractérisée en ce que la partie de profil aérodynamique (2) comprend un agencement de refroidissement ayant un système d'admission (28) permettant d'introduire un liquide de refroidissement (26) dans le profil aérodynamique (2) pour refroidir le profil aérodynamique (2), au moins un premier passage de refroidissement (20, 40, 68) permettant d'amener le liquide de refroidissement dans une direction allant du bord d'attaque (5) vers le bord de fuite (4) sur le côté d'aspiration (7), dans laquelle le premier passage de refroidissement (20, 40, 68) est connecté sur le plan fluide au système d'admission (28), au moins un second passage de refroidissement (31, 50, 69) permettant d'amener le liquide de refroidissement dans une direction allant du bord d'attaque (4) vers le bord de fuite (5) sur le côté de pression (6), dans laquelle le second passage de refroidissement (31, 50, 69) est connecté sur le plan liquide au système d'admission (28) et dans laquelle le liquide de refroidissement (26) s'écoulant dans le second passage de refroidissement (31, 50, 69) est ressorti de la partie de profil aérodynamique (2) à travers une ouverture (13) pratiquée au niveau du bord de fuite (5).
PCT/EP2013/077812 2013-01-03 2013-12-20 Aube de turbomachine WO2014106598A1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
EP13150152.0 2013-01-03
EP13150152.0A EP2752554A1 (fr) 2013-01-03 2013-01-03 Pale pour turbomachine

Publications (1)

Publication Number Publication Date
WO2014106598A1 true WO2014106598A1 (fr) 2014-07-10

Family

ID=47623913

Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/EP2013/077812 WO2014106598A1 (fr) 2013-01-03 2013-12-20 Aube de turbomachine

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EP (1) EP2752554A1 (fr)
WO (1) WO2014106598A1 (fr)

Families Citing this family (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US10370976B2 (en) 2017-08-17 2019-08-06 United Technologies Corporation Directional cooling arrangement for airfoils
GB201720829D0 (en) * 2017-12-14 2018-01-31 Rolls Royce Plc Aerofoil and method of manufacture
US10570748B2 (en) 2018-01-10 2020-02-25 United Technologies Corporation Impingement cooling arrangement for airfoils
US10787932B2 (en) * 2018-07-13 2020-09-29 Honeywell International Inc. Turbine blade with dust tolerant cooling system

Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5813835A (en) * 1991-08-19 1998-09-29 The United States Of America As Represented By The Secretary Of The Air Force Air-cooled turbine blade
EP1881157A1 (fr) * 2006-07-18 2008-01-23 United Technologies Corporation Microcircuits en serpentin pour un enlèvement local de chaleur
EP1936118A2 (fr) * 2006-12-11 2008-06-25 United Technologies Corporation Modifications du noyau de coulée des aubes de turbine pour microcircuits à serpentin périphérique

Patent Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5813835A (en) * 1991-08-19 1998-09-29 The United States Of America As Represented By The Secretary Of The Air Force Air-cooled turbine blade
EP1881157A1 (fr) * 2006-07-18 2008-01-23 United Technologies Corporation Microcircuits en serpentin pour un enlèvement local de chaleur
EP1936118A2 (fr) * 2006-12-11 2008-06-25 United Technologies Corporation Modifications du noyau de coulée des aubes de turbine pour microcircuits à serpentin périphérique

Also Published As

Publication number Publication date
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