EP2087206B1 - Aube de turbine - Google Patents

Aube de turbine Download PDF

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Publication number
EP2087206B1
EP2087206B1 EP07820379A EP07820379A EP2087206B1 EP 2087206 B1 EP2087206 B1 EP 2087206B1 EP 07820379 A EP07820379 A EP 07820379A EP 07820379 A EP07820379 A EP 07820379A EP 2087206 B1 EP2087206 B1 EP 2087206B1
Authority
EP
European Patent Office
Prior art keywords
cooling
turbine blade
elements
wall section
leading edge
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Not-in-force
Application number
EP07820379A
Other languages
German (de)
English (en)
Other versions
EP2087206A1 (fr
Inventor
Heinz-Jürgen GROSS
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Siemens AG
Original Assignee
Siemens AG
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 AG filed Critical Siemens AG
Priority to EP07820379A priority Critical patent/EP2087206B1/fr
Publication of EP2087206A1 publication Critical patent/EP2087206A1/fr
Application granted granted Critical
Publication of EP2087206B1 publication Critical patent/EP2087206B1/fr
Not-in-force legal-status Critical Current
Anticipated expiration legal-status Critical

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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
    • 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
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/201—Heat transfer, e.g. cooling by impingement of a fluid
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/221—Improvement of heat transfer
    • F05D2260/2214—Improvement of heat transfer by increasing the heat transfer surface
    • F05D2260/22141—Improvement of heat transfer by increasing the heat transfer surface using fins or ribs

Definitions

  • the invention relates to a turbine blade according to the preamble of claim 1.
  • Turbine blades in particular turbine blades for gas turbines are exposed to high temperatures during operation, which quickly exceed the limit of the material stress. This applies in particular to the regions in the vicinity of the flow inlet edge at which the hot process gas flow first of all occurs on the blade profile of the turbine blade.
  • turbine blades In order to use turbine blades even at high temperatures, it has long been known to cool turbine blades suitable, so that they have a higher temperature resistance. With turbine blades, which have a higher temperature resistance, higher energy efficiencies can be achieved in particular.
  • Cooling types include convection cooling, impingement cooling and film cooling.
  • convection cooling cooling air passes through channels in the interior of the blade and uses the convective effect to dissipate the heat.
  • impingement cooling a cooling air flow impinges on the inner blade surface from the inside. In this way, a very good cooling effect is made possible at the point of impact, but this is limited only to the narrow area of the point of impact and the closer environment.
  • This type of cooling is therefore usually used for cooling the flow inlet edge, which is also referred to as the leading edge, a turbine blade.
  • film cooling cooling air is directed out through openings in the turbine blade from the interior of the turbine blade. This cooling air flows around the turbine blade and forms an insulating layer between the hot process gas and the blade surface.
  • the described Cooling types are suitably combined depending on the application, in order to achieve the most effective blade cooling possible.
  • coolants such as turbulators, which are mostly provided in the form of ribs
  • turbulators which are mostly provided in the form of ribs
  • These are arranged within the cooling channels provided for the convection flow, which run in the interior of the turbine blade.
  • the incorporation of fins in the cooling channels causes the flow of cooling air in the boundary layers to be detached and entangled. Due to the forced disruption of the flow, the heat transfer can be increased in the presence of a temperature difference between the cooling channel wall and the cooling air.
  • the ribbing constantly causes the flow to form new "recovery areas" in which a substantial increase in the local heat transfer coefficient can be achieved.
  • turbine blades In order to cool the flow inlet leading edge, ie leading edge, of turbine blades, which is usually very heavily stressed during operation, turbine blades often run parallel and close to the flow inlet edge Cooling channels formed, which is supplied by further formed in the blades cooling channels cooling air.
  • the convective cooling of the flow inlet edge realized in this way is usually supplemented by impingement cooling of the inner wall of the cooling channel extending near the flow inlet edge in the case of film-sensed blades.
  • convective cooling is intensified by turbulators disposed on the inner wall of the cooling duct.
  • the invention has for its object to provide a turbine blade, which can be cooled more effectively compared to known solutions both existing and in the absence of film cooling and has a longer service life.
  • the turbine blade has a front edge extending on one side of the turbine blade, the cooling channel being delimited by a wall portion opposite the leading edge and having two or more peg-shaped cooling elements of different lengths extending into the cooling channel from this wall portion and the length thereof is different for adaptation to the local cooling demand.
  • the generally thermally stressed front edge can thus be cooled very effectively.
  • the cooling elements according to the invention which extend from the wall portion into the cooling channel, and in particular cause a strong turbulence of the cooling medium, the heat transfer can be significantly increased at a temperature difference between the wall portion and the cooling medium, along with a substantial increase of the local heat transfer coefficients. Overall, in this way, the heat in the vicinity of the leading edge can be dissipated very effectively, along with a very effective cooling of the leading edge.
  • the cooling elements first impinged by the cooling medium in a bump-cooling manner are in the form of pegs.
  • Pine-shaped cooling elements cause on the one hand an enlargement of the coolable wall surface and on the other hand after impact cooling a very strong turbulence of the cooling medium, for example in the form of cooling air, which is enforced by the strong forced disruption of the flow at a temperature difference between a wall of the cooling channel and the cooling medium Heat transfer can be increased, along with a significant increase in the local heat transfer coefficient.
  • the invention provided peg-shaped design of the cooling elements which are formed during operation of the turbine blade in the cooling elements thermal stresses are minimized, so that it can come to any êtanrissen, in particular in this case the thermal stresses are significantly lower than the thermal stresses form in known turbulators. According to the invention, therefore, the entire voltage situation is improved and it can be a significant increase in the life of the cooling elements over known solutions can be achieved, with the high life of the cooling elements and a long service life and service life of the turbine blade is connected.
  • the turbine blade according to the invention can be exposed to higher gas temperatures compared to known solutions, even if no film cooling is provided. If film cooling is provided, even higher gas temperatures are possible. This in turn gives rise to the possibility of being able to form the turbine blade according to the invention with thinner outer walls.
  • the cooling capacity of each individual peg-shaped cooling element over a suitably formed length is equal to the predetermined local cooling requirement in the environment of the Cooling element adapted. Cooling elements in the vicinity of which a high cooling requirement exists, according to the invention have a greater length than cooling elements in the environment of the cooling demand is less pronounced. By increasing the length of a single cooling element, on the one hand, the "swirl area" and the surface to be cooled are increased, along with a significant increase in the local heat transfer coefficient.
  • the wall section has a wall surface facing the cooling channel, wherein the at least one cooling element or the two or more Cooling elements extend orthogonal to the wall surface or orthogonal to the curved wall surface in the cooling channel inside.
  • the inventively provided extension in a direction orthogonal to the wall surface of the cooling channel causes a very effective turbulence of the cooling medium, which is accompanied by a very effective cooling, in particular the leading edge, since according to the invention a directed substantially perpendicular to the longitudinal extent of the cooling elements directed flow of the cooling elements with the cooling medium can.
  • the cooling channel is preferably limited by a wall portion facing the cooling channel having a curved wall surface, wherein two or more cooling elements are provided, wherein the cooling elements have a longitudinal extent extending into the cooling channel, and wherein the two or more cooling elements are directed with their longitudinal extent to the center of the curvature of the wall surface.
  • cooling elements which are directed with their longitudinal extent to the center of the curvature of the wall surface, a very effective turbulence of the cooling elements flowing against the cooling medium can be achieved.
  • the convection cooling realized by means of the cooling elements can be very effectively combined with impingement cooling, such that the cooling medium flows in such a way onto the cooling elements that it impinges on the cooling elements, so that a very high cooling effect is achieved in the respective impingement point can be, which causes in conjunction with the provided convection cooling a very effective cooling of the turbine blade according to the invention.
  • turbine blades In operation, turbine blades generally have a very inhomogeneous temperature distribution, which is associated with large thermal loads acting on the turbine blades, which in particular have a detrimental effect on the service life of the turbine blade.
  • an inhomogeneous temperature distribution forming along the radial direction results for the leading edge.
  • the temperature distribution for example at the leading edge, can be "evened out", since according to the invention in comparatively hot places by appropriately trained cooling elements, a correspondingly strong cooling and vice versa.
  • the turbine blade according to the invention can thus be cooled in a manner which counteracts an inhomogeneous temperature distribution, which is advantageous in particular with regard to effective cooling of the leading edge.
  • a cooling channel partially delimiting the wall portion opposite rear wall is provided in which one or more impingement cooling holes are provided.
  • impingement cooling holes are provided. These are preferably placed and aligned in the back wall so that the cooling air jets flowing through them are directed onto the cooling elements, whereby a particularly efficient cooling of the leading edge can be achieved.
  • the distance between the cooling element tip, on the one hand, and the mouth of the impact cooling opening, on the other hand can be kept comparatively small.
  • This also applies to a comparatively large outflow cross section of the cooling channel. A disturbance of the impingement cooling jets by transversely to the rays, d. H. along the cooling channel flowing cooling air can thus be safely avoided.
  • the invention relates generally to a turbine blade having a leading edge, a cooling channel formed in the turbine blade for carrying cooling air, which extends at least partially along the leading edge, and a number of cooling elements, which are arranged in the longitudinal direction of the cooling channel in this sequentially stationary, each individual cooling element has a cooling capacity which is adapted to a predetermined cooling requirement for the leading edge in the vicinity of the cooling element, and wherein the cooling channel preferably extends parallel to the leading edge through the turbine blade.
  • FIG. 1 shows a sketch-like sectional view of a front portion of an airfoil of a turbine blade 10 according to the invention, with a flat sectional surface perpendicular to the front edge 12.
  • the leading edge 12 may also be referred to as a flow inlet edge.
  • a cooling channel 14 extending parallel to the front edge 12 (ie, a radially extending channel 14 in the case of axially through-flowed turbines) is formed near the front edge 12, which is delimited by a wall section 24 in relation to the front edge 12.
  • peg-shaped cooling elements 18 extend into the cooling channel 14, wherein the cooling elements 18 are directed with their longitudinal extent to the center of the curvature of the wall surface 16.
  • openings 22 are formed to supply the cooling channel 14 of further cooling channels (not shown), which are formed in the rear region of the turbine blade 10, cooling air chilling cooling.
  • FIG. 2 shows a further sectional view of the front portion of the turbine blade 10 according to the invention, with a flat sectional surface parallel to the leading edge 12.
  • the formed on the curved wall surface 16 of the cooling channel 14 cooling elements 18 extend orthogonally from the curved wall surface 16 into the cooling channel 14. As out FIG. 2 in the radial direction R, the length of the cooling elements 18 varies. According to the invention, this counteracts the inhomogeneous temperature distribution which forms along the leading edge 12 when the turbine blade 10 is used.
  • the frusto-conical cooling elements 18 have a greater length in the middle region than in the edge regions, since, as stated above, by increasing the length of the cooling elements 18, the local heat transfer coefficient and thus the cooling capacity the cooling elements 18 can be increased.
  • the impingement cooling comprises the impact of cooling air emerging from the openings 22 on the arched wall surface 16 or the cooling elements 18 in order to locally enable a very good cooling effect there.
  • the invention provides that the cooling elements 18 are directed with their longitudinal extent to the center of the curvature of the wall surface 16, a very effective impingement cooling can be provided with the overall convection cooling in conjunction with a very effective cooling of the turbine blade 10 can be provided ,
  • the cooling passage 14 is opened on both sides of the turbine blade 10 to flow the cooling air in two directions out of the cooling passage 14. As a result, a temperature harmonization of the turbine blade 10 is favored, since where cooling air is needed, cooling air is provided, and the effect of the impingement cooling is not reduced by a cross-flow.
  • the cooling elements 18 may also be formed rib-shaped, which extend along the cooling channel 14, ie in the flow direction of the cooling air.
  • the surface of the wall surface 16 is significantly increased in order to improve the cooling of the then preferably convectively cooled turbine blade 10. It is conceivable that the height of the ribs due to the aforementioned locally different temperatures at the front edge 12 can be adapted to match.

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

Claims (7)

  1. Aube ( 10 ) de turbine,
    comprenant une lame d'aube ayant un canal ( 14 ) de refroidissement et un bord ( 12 ) avant s'étendant le long de la lame de l'aube, le canal ( 14 ) de refroidissement étant délimité par rapport au bord ( 12 ) avant par une section ( 24 ) de paroi, des moyens étant prévus pour refroidir la section ( 24 ) de paroi par rebondissement,
    caractérisée en ce que
    en partant de la section ( 24 ) de paroi deux ou plusieurs éléments ( 18 ) de refroidissement en forme de tenon s'étendent dans le canal ( 14 ) de refroidissement en ayant des longueurs différentes, leurs longueurs étant adaptées à un besoin local prescrit de refroidissement.
  2. Aube ( 10 ) de turbine suivant la revendication 1,
    dans laquelle la section ( 24 ) de paroi a une surface ( 16 ) de paroi tournée vers le canal ( 14 ) de refroidissement et dans laquelle le au moins un élément ( 18 ) de refroidissement s'étend dans le canal ( 14 ) de refroidissement orthogonalement à la surface ( 16 ) de paroi.
  3. Aube ( 10 ) de turbine suivant la revendication 1 ou 2, dans laquelle la section ( 24 ) de paroi a une surface ( 16 ) de paroi courbée et tournée vers le canal ( 14 ) de refroidissement, dans laquelle il est prévu deux ou plusieurs éléments ( 18 ) de refroidissement, dans laquelle les éléments ( 18 ) de refroidissement ont une étendue en longueur s'étendant à l'intérieur du canal ( 14 ) de refroidissement et dans laquelle deux ou plusieurs éléments ( 18 ) de refroidissement sont dirigés par leur étendue en longueur sur le centre de la courbure de la surface ( 16 ) de paroi.
  4. Aube ( 10 ) de turbine suivant l'une des revendications précédentes,
    dans laquelle le au moins un élément ( 18 ) de refroidissement ou deux ou plusieurs éléments ( 18 ) de refroidissement est ou sont d'une seule pièce avec la section ( 24 ) de paroi.
  5. Aube ( 10 ) de turbine suivant l'une des revendications précédentes,
    dans laquelle le canal ( 14 ) de refroidissement s'étend au moins en partie parallèlement au bord ( 12 ) avant en passant à travers l'aube ( 10 ) de turbine.
  6. Aube ( 10 ) de turbine suivant l'une des revendications précédentes,
    dans laquelle le moyen de refroidissement par rebondissement de la section ( 24 ) de paroi est une paroi ( 20 ) arrière qui délimite le canal ( 14 ) de refroidissement, qui est opposée à la section ( 24 ) de paroi et dans laquelle plusieurs ouvertures ( 22 ) de refroidissement par rebondissement sont prévues.
  7. Aube ( 10 ) de turbine suivant la revendication 6,
    dans laquelle les ouvertures ( 22 ) de refroidissement par rebondissement sont disposées de façon à ce que les faisceaux d'air de refroidissement y passant soient dirigés sur les éléments ( 18 ) de refroidissement.
EP07820379A 2006-11-08 2007-09-20 Aube de turbine Not-in-force EP2087206B1 (fr)

Priority Applications (1)

Application Number Priority Date Filing Date Title
EP07820379A EP2087206B1 (fr) 2006-11-08 2007-09-20 Aube de turbine

Applications Claiming Priority (3)

Application Number Priority Date Filing Date Title
EP06023274A EP1921268A1 (fr) 2006-11-08 2006-11-08 Aube de turbine
EP07820379A EP2087206B1 (fr) 2006-11-08 2007-09-20 Aube de turbine
PCT/EP2007/059935 WO2008055737A1 (fr) 2006-11-08 2007-09-20 Aube de turbine

Publications (2)

Publication Number Publication Date
EP2087206A1 EP2087206A1 (fr) 2009-08-12
EP2087206B1 true EP2087206B1 (fr) 2010-03-03

Family

ID=37951488

Family Applications (2)

Application Number Title Priority Date Filing Date
EP06023274A Withdrawn EP1921268A1 (fr) 2006-11-08 2006-11-08 Aube de turbine
EP07820379A Not-in-force EP2087206B1 (fr) 2006-11-08 2007-09-20 Aube de turbine

Family Applications Before (1)

Application Number Title Priority Date Filing Date
EP06023274A Withdrawn EP1921268A1 (fr) 2006-11-08 2006-11-08 Aube de turbine

Country Status (7)

Country Link
US (1) US8297926B2 (fr)
EP (2) EP1921268A1 (fr)
JP (2) JP2010509532A (fr)
CN (1) CN101535602B (fr)
AT (1) ATE459785T1 (fr)
DE (1) DE502007003044D1 (fr)
WO (1) WO2008055737A1 (fr)

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US9957816B2 (en) 2014-05-29 2018-05-01 General Electric Company Angled impingement insert
US9995148B2 (en) 2012-10-04 2018-06-12 General Electric Company Method and apparatus for cooling gas turbine and rotor blades
US10690055B2 (en) 2014-05-29 2020-06-23 General Electric Company Engine components with impingement cooling features

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KR101513474B1 (ko) * 2013-02-27 2015-04-23 두산중공업 주식회사 터빈 블레이드
US9850762B2 (en) 2013-03-13 2017-12-26 General Electric Company Dust mitigation for turbine blade tip turns
US20150204197A1 (en) * 2014-01-23 2015-07-23 Siemens Aktiengesellschaft Airfoil leading edge chamber cooling with angled impingement
US10001013B2 (en) * 2014-03-06 2018-06-19 General Electric Company Turbine rotor blades with platform cooling arrangements
US10364684B2 (en) 2014-05-29 2019-07-30 General Electric Company Fastback vorticor pin
US10422235B2 (en) 2014-05-29 2019-09-24 General Electric Company Angled impingement inserts with cooling features
EP3149279A1 (fr) 2014-05-29 2017-04-05 General Electric Company Générateur de turbulence fastback
WO2016007145A1 (fr) * 2014-07-09 2016-01-14 Siemens Aktiengesellschaft Système de canaux d'amorçage de jets d'impact à l'intérieur de systèmes de refroidissement
US10280785B2 (en) 2014-10-31 2019-05-07 General Electric Company Shroud assembly for a turbine engine
US10233775B2 (en) 2014-10-31 2019-03-19 General Electric Company Engine component for a gas turbine engine
US20160201476A1 (en) * 2014-10-31 2016-07-14 General Electric Company Airfoil for a turbine engine
US20160333701A1 (en) * 2015-05-12 2016-11-17 United Technologies Corporation Airfoil impingement cavity
US20170107827A1 (en) * 2015-10-15 2017-04-20 General Electric Company Turbine blade
US10352177B2 (en) 2016-02-16 2019-07-16 General Electric Company Airfoil having impingement openings
KR101906701B1 (ko) * 2017-01-03 2018-10-10 두산중공업 주식회사 가스터빈 블레이드
EP3396297A1 (fr) * 2017-04-28 2018-10-31 Siemens Aktiengesellschaft Dispositif de refroidissement
US10830049B2 (en) 2017-05-02 2020-11-10 Raytheon Technologies Corporation Leading edge hybrid cavities and cores for airfoils of gas turbine engine
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Publication number Priority date Publication date Assignee Title
US9995148B2 (en) 2012-10-04 2018-06-12 General Electric Company Method and apparatus for cooling gas turbine and rotor blades
US9957816B2 (en) 2014-05-29 2018-05-01 General Electric Company Angled impingement insert
US10690055B2 (en) 2014-05-29 2020-06-23 General Electric Company Engine components with impingement cooling features

Also Published As

Publication number Publication date
DE502007003044D1 (de) 2010-04-15
JP2012137089A (ja) 2012-07-19
EP2087206A1 (fr) 2009-08-12
CN101535602A (zh) 2009-09-16
ATE459785T1 (de) 2010-03-15
CN101535602B (zh) 2012-01-11
WO2008055737A1 (fr) 2008-05-15
EP1921268A1 (fr) 2008-05-14
US20100143153A1 (en) 2010-06-10
JP5269223B2 (ja) 2013-08-21
US8297926B2 (en) 2012-10-30
JP2010509532A (ja) 2010-03-25

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