EP1496203B1 - Aube de turbine à gaz avec refroidissement par impact - Google Patents

Aube de turbine à gaz avec refroidissement par impact Download PDF

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Publication number
EP1496203B1
EP1496203B1 EP04090262A EP04090262A EP1496203B1 EP 1496203 B1 EP1496203 B1 EP 1496203B1 EP 04090262 A EP04090262 A EP 04090262A EP 04090262 A EP04090262 A EP 04090262A EP 1496203 B1 EP1496203 B1 EP 1496203B1
Authority
EP
European Patent Office
Prior art keywords
cooling
air
impingement
wall
turbine blade
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.)
Expired - Fee Related
Application number
EP04090262A
Other languages
German (de)
English (en)
Other versions
EP1496203A1 (fr
Inventor
Peter Davison
Barbara Blume
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.)
Rolls Royce Deutschland Ltd and Co KG
Original Assignee
Rolls Royce Deutschland Ltd and Co KG
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 Rolls Royce Deutschland Ltd and Co KG filed Critical Rolls Royce Deutschland Ltd and Co KG
Publication of EP1496203A1 publication Critical patent/EP1496203A1/fr
Application granted granted Critical
Publication of EP1496203B1 publication Critical patent/EP1496203B1/fr
Anticipated expiration legal-status Critical
Expired - Fee Related legal-status Critical Current

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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
    • 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
    • F05D2250/00—Geometry
    • F05D2250/10—Two-dimensional
    • F05D2250/14—Two-dimensional elliptical
    • 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
    • F05D2250/712—Shape curved concave
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/201—Heat transfer, e.g. cooling by impingement of a fluid

Definitions

  • the invention relates to a turbine blade with an impingement cooling of the thermally highly loaded outer wall sections, wherein inside the hollow turbine blade at least one partition to form a cooling air supplied cooling air chamber and in the partition a plurality of impeller air ducts for applying the formation of an impingement air cooling chamber at a distance adjacent inner surface of the hot Outside wall sections is provided with cooling air.
  • the efficiency of gas turbines can be improved by increasing the combustion temperatures achieved in the combustion chamber.
  • a temperature increase is limited insofar as the thermal capacity of the components exposed to the hot gases, in particular the guide vanes and rotor blades which are also subject to high mechanical loads, is limited in the turbine stage connected downstream of the combustion chamber.
  • the relevant components and in particular their thermally highly stressed areas are known to be cooled with branched off from the compressor cooling air.
  • an impingement cooling system for a turbine blade known, for example, from EP 1 001 135 A2
  • longitudinally extending partitions are arranged in the interior of a hollow blade delimited by two side walls, each one having an elongate side wall portion Forming cooling air supply and distribution chamber (cooling air chamber) and a plurality of these adjacent impact air cooling chambers.
  • cooling air chamber Cooling air chamber
  • the cooling air introduced into the cooling air chamber passes successively - in other cases also simultaneously - into the adjacent impingement air cooling chambers, thereby intensively cooling the inner surfaces of the thermally highly stressed areas of the outer walls of the turbine blade from inside, and thus with the gas turbine at the highest possible combustion temperatures high efficiency and to be able to operate without material damage.
  • the impingement air ducts are rectilinear in the dividing wall but obliquely oriented to provide a favorable angle for impingement of the impingement cooling air on the inner surfaces of the outer walls.
  • the air emerging from the impingement air cooling chambers via air channels in the side walls of the turbine blade also creates an insulating layer between the blade material and the hot gas, which further reduces the thermal load on the turbine blade.
  • the impact air ducts reduce the load-bearing surface of the partitions which carry the outer walls, and on the other hand tension peaks associated with high local mechanical stress occur in the area of the impingement air ducts, resulting in a reduction in the service life of the turbine blade.
  • the strength of the partitions which would lead to a reduction of the local voltage peaks with correspondingly large dimensions, not be increased arbitrarily.
  • the invention has the object of providing a turbine blade of the type mentioned in such a way that at substantially unchanged weight, the voltage spikes degraded in the area of the baffles and thus the Zeitschwing- and creep resistance and ultimately the life can be increased.
  • the object is achieved with a trained according to the features of claim 1 turbine blade. From the dependent claim, there are further features of the invention.
  • the invention is based on the recognition that the partitions in the middle region are the coolest and represent a region of highest tensile stress.
  • this area are in the turbine blades formed according to the prior art, the inlet openings of the rectilinear and to achieve a certain air impact angle obliquely aligned impingement air ducts, so that there the concentration of stress is particularly high.
  • the impingement air ducts are now bent, in such a way that the impact air outlet location and angle remains unchanged and the impingement air is directed at a predetermined angle to the inner surface of the relevant outer wall section, but the air inlet opening and thus the entire impingement air duct in a warmer Edge zone of the partition is laid with lower tensile stresses.
  • the impingement air channel is concavely curved with respect to the outer wall and runs as a whole in the vicinity of the hot outer wall and quasi parallel to this.
  • This formation and arrangement of the baffles reduces the notch effect and increases the creep strength and the time fatigue strength, so that the life of the turbine blade is increased.
  • the reduction of the stress concentration in the area of the baffles thus achieved allows for smaller wall thicknesses of the baffles, so that the weight of the turbine blade can be reduced.
  • the cross-sectional area of the baffles has the shape of a slot or oval, wherein the longitudinal axis of the oval or elongated hole extends in the longitudinal direction of the cooling air chamber.
  • This cross-sectional shape and its radial orientation and the consequent low notch factor creep and time-swinging behavior is also improved and increases the life of the turbine blade.
  • the wall thickness of the partitions can be reduced and thus the weight of the turbine blade can be reduced.
  • the blade profile 1 of a high-pressure turbine blade is formed from a thin-walled outer wall 2 and supporting inner partitions 3 to 5.
  • the impingement air cooling chamber 8 is defined by the first partition wall 3 and an outer wall portion 2b, and the second impingement air cooling chamber 9 is formed by the second partition wall 4, two outer wall portions 2c, 2d, and the third partition wall 5.
  • the third partition wall 5 and two outer wall sections 2e, 2f include another cooling chamber 10.
  • the cooling air supplied to the cooling air chamber 6 flows via the due to their curvature throughout in a hot, relatively low-tension region near the outer wall 2 in the first and second partition wall 3, 4 extending impingement air ducts 7 in the first and in the second impingement air cooling chamber 8 and 9, in the cooling air impinges on the inner surfaces of the adjacent outer wall sections 2 b and 2 c and 2 d, thereby intensively cooling them.
  • the cooling air introduced into the first impact air cooling chamber 8 passes via air ducts 11a in the outer wall section 2b to the outer surface, in order there to form an air layer for external shielding of the material with respect to the hot air.
  • the cooling air in the second impact air cooling chamber 9 flows outward via the cooling chamber 10 and cooling passages 11b or directly via the cooling passages 11c.
  • the cross-sectional area of the Impeller air channels 7, as shown in FIG. 2 has the shape of a slot and the longitudinal axis of the cross-sectional area coincides with the longitudinal axis of the blade profile 1 or its radial orientation.
  • the cross-sectional area of the impingement air ducts may equally be that of an ellipse.
  • the elliptical or oblong-shaped design of the impingement air ducts in conjunction with the orientation of the longitudinal axis of the cross-sectional area to the dominant load vector, on the one hand increases the time fatigue strength and, on the other hand, reduces the notch effect, so that a longer service life of the high-pressure turbine blade can be achieved.

Landscapes

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

Claims (2)

  1. Aube de turbine avec refroidissement par impact des sections de paroi extérieure soumises à une forte contrainte thermique, sachant qu'à l'intérieur de l'aube de turbine creuse sont prévus au moins une cloison pour former une chambre d'air de refroidissement alimentée en air de refroidissement et dans la cloison plusieurs canaux d'air d'impact destinés à alimenter en air de refroidissement d'impact la surface intérieure de la (des) section(s) de paroi extérieure chaude(s) voisine à une certaine distance par formation d'une chambre de refroidissement à air d'impact, caractérisée en ce que les canaux d'air d'impact (7) sont incurvés de manière concave par rapport à la paroi extérieure proche (2) et pour l'essentiel parallèles à celle-ci et entièrement disposés dans la zone chaude proche de la paroi extérieure.
  2. Aube de turbine selon la revendication nº 1, caractérisée en ce que les canaux d'air d'impact (7) présentent une surface de section transversale de forme oblongue ou elliptique, avec un axe longitudinal dont l'orientation correspond à l'orientation radiale de l'aube.
EP04090262A 2003-07-11 2004-06-28 Aube de turbine à gaz avec refroidissement par impact Expired - Fee Related EP1496203B1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
DE10332563 2003-07-11
DE10332563A DE10332563A1 (de) 2003-07-11 2003-07-11 Turbinenschaufel mit Prallkühlung

Publications (2)

Publication Number Publication Date
EP1496203A1 EP1496203A1 (fr) 2005-01-12
EP1496203B1 true EP1496203B1 (fr) 2006-02-08

Family

ID=33441771

Family Applications (1)

Application Number Title Priority Date Filing Date
EP04090262A Expired - Fee Related EP1496203B1 (fr) 2003-07-11 2004-06-28 Aube de turbine à gaz avec refroidissement par impact

Country Status (3)

Country Link
US (1) US7063506B2 (fr)
EP (1) EP1496203B1 (fr)
DE (2) DE10332563A1 (fr)

Families Citing this family (22)

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Publication number Priority date Publication date Assignee Title
US20050265840A1 (en) * 2004-05-27 2005-12-01 Levine Jeffrey R Cooled rotor blade with leading edge impingement cooling
US7217094B2 (en) * 2004-10-18 2007-05-15 United Technologies Corporation Airfoil with large fillet and micro-circuit cooling
GB2420156B (en) 2004-11-16 2007-01-24 Rolls Royce Plc A heat transfer arrangement
US8966754B2 (en) * 2006-11-21 2015-03-03 General Electric Company Methods for reducing stress on composite structures
US7704048B2 (en) * 2006-12-15 2010-04-27 Siemens Energy, Inc. Turbine airfoil with controlled area cooling arrangement
GB0811391D0 (en) * 2008-06-23 2008-07-30 Rolls Royce Plc A rotor blade
EP2196625A1 (fr) * 2008-12-10 2010-06-16 Siemens Aktiengesellschaft Aube de turbine dotée d'un passage agencé dans une paroi de séparation et noyau de coulage associé
US20110110790A1 (en) * 2009-11-10 2011-05-12 General Electric Company Heat shield
US9347324B2 (en) 2010-09-20 2016-05-24 Siemens Aktiengesellschaft Turbine airfoil vane with an impingement insert having a plurality of impingement nozzles
US8840370B2 (en) 2011-11-04 2014-09-23 General Electric Company Bucket assembly for turbine system
US9004866B2 (en) * 2011-12-06 2015-04-14 Siemens Aktiengesellschaft Turbine blade incorporating trailing edge cooling design
EP2828484B1 (fr) * 2012-03-22 2019-05-08 Ansaldo Energia IP UK Limited Aube de turbine
US9506351B2 (en) * 2012-04-27 2016-11-29 General Electric Company Durable turbine vane
EP2895696A1 (fr) * 2012-08-06 2015-07-22 General Electric Company Composant de turbine rotatif à alignement de trou préférentiel
US9394798B2 (en) 2013-04-02 2016-07-19 Honeywell International Inc. Gas turbine engines with turbine airfoil cooling
US10145246B2 (en) 2014-09-04 2018-12-04 United Technologies Corporation Staggered crossovers for airfoils
EP3000970B1 (fr) * 2014-09-26 2019-06-12 Ansaldo Energia Switzerland AG Système de refroidissement pour le bord d'attaque d'une aube de turbine d'une turbine à gaz
US10208603B2 (en) 2014-11-18 2019-02-19 United Technologies Corporation Staggered crossovers for airfoils
US10633980B2 (en) 2017-10-03 2020-04-28 United Technologies Coproration Airfoil having internal hybrid cooling cavities
US10626734B2 (en) 2017-10-03 2020-04-21 United Technologies Corporation Airfoil having internal hybrid cooling cavities
US10626733B2 (en) 2017-10-03 2020-04-21 United Technologies Corporation Airfoil having internal hybrid cooling cavities
US10704398B2 (en) * 2017-10-03 2020-07-07 Raytheon Technologies Corporation Airfoil having internal hybrid cooling cavities

Family Cites Families (9)

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Publication number Priority date Publication date Assignee Title
US5674050A (en) * 1988-12-05 1997-10-07 United Technologies Corp. Turbine blade
US5660524A (en) * 1992-07-13 1997-08-26 General Electric Company Airfoil blade having a serpentine cooling circuit and impingement cooling
US5403158A (en) 1993-12-23 1995-04-04 United Technologies Corporation Aerodynamic tip sealing for rotor blades
DE19848104A1 (de) * 1998-10-19 2000-04-20 Asea Brown Boveri Turbinenschaufel
US6036441A (en) * 1998-11-16 2000-03-14 General Electric Company Series impingement cooled airfoil
JP2000265802A (ja) * 1999-01-25 2000-09-26 General Electric Co <Ge> ガスタービン動翼冷却通路連絡路
US6206638B1 (en) * 1999-02-12 2001-03-27 General Electric Company Low cost airfoil cooling circuit with sidewall impingement cooling chambers
DE10059997B4 (de) 2000-12-02 2014-09-11 Alstom Technology Ltd. Kühlbare Schaufel für eine Gasturbinenkomponente
FR2829174B1 (fr) * 2001-08-28 2006-01-20 Snecma Moteurs Perfectionnement apportes aux circuits de refroidissement pour aube de turbine a gaz

Also Published As

Publication number Publication date
US7063506B2 (en) 2006-06-20
EP1496203A1 (fr) 2005-01-12
DE502004000285D1 (de) 2006-04-20
US20050111981A1 (en) 2005-05-26
DE10332563A1 (de) 2005-01-27

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