EP1921269A1 - Turbinenschaufel - Google Patents

Turbinenschaufel Download PDF

Info

Publication number
EP1921269A1
EP1921269A1 EP06023377A EP06023377A EP1921269A1 EP 1921269 A1 EP1921269 A1 EP 1921269A1 EP 06023377 A EP06023377 A EP 06023377A EP 06023377 A EP06023377 A EP 06023377A EP 1921269 A1 EP1921269 A1 EP 1921269A1
Authority
EP
European Patent Office
Prior art keywords
cooling
ribs
turbine blade
rib
pair
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.)
Withdrawn
Application number
EP06023377A
Other languages
German (de)
English (en)
French (fr)
Inventor
Heinz-Jürgen Dr. 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
Siemens Corp
Original Assignee
Siemens AG
Siemens Corp
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, Siemens Corp filed Critical Siemens AG
Priority to EP06023377A priority Critical patent/EP1921269A1/de
Priority to US12/513,682 priority patent/US8215909B2/en
Priority to JP2009535661A priority patent/JP5329418B2/ja
Priority to PCT/EP2007/061127 priority patent/WO2008055764A1/de
Priority to EP07821492.1A priority patent/EP2087207B1/de
Publication of EP1921269A1 publication Critical patent/EP1921269A1/de
Withdrawn legal-status Critical Current

Links

Images

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
    • 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
    • F01D5/147—Construction, i.e. structural features, e.g. of weight-saving hollow blades
    • 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
    • F05D2250/18—Two-dimensional patterned
    • F05D2250/181—Two-dimensional patterned ridged
    • 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
    • F05D2250/28—Three-dimensional patterned
    • 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
    • F05D2250/34—Arrangement of components translated
    • 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
    • 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.
  • Turbine blades particularly turbine blades for gas turbines, are exposed during operation to high temperatures which rapidly exceed the limit of material stress. This applies in particular to the areas in the vicinity of the flow inlet edge.
  • 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 cooling is probably the most common type of blade cooling.
  • This type of cooling cooling air is passed 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 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 of a turbine blade, which is exposed to high temperature loads.
  • 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 types of cooling described are suitably combined depending on the application in order to achieve the most effective blade cooling possible.
  • coolants such as turbulators, which are usually provided in the form of ribs
  • turbulators which are usually 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.
  • cooling channels are often formed in turbine blades parallel to and close to the flow inlet edge, to which cooling air is supplied by further cooling channels formed in the blades.
  • 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 whose flow inlet edge can be cooled compared to known solutions more effectively, both in existing as well as non-existing film cooling.
  • a turbine blade which has a plurality of ribs, which are arranged successively in a cooling channel which extends along a flow inlet edge, and in which each with two ribs a pair of ribs is formed, arranged the ribs in skating step shape are.
  • the inventively provided pairwise arrangement of the ribs in skate step shape causes over known solutions a greatly increased turbulence of the cooling air flowing in the cooling channel according to the invention, such that the cooling air flowing in the cooling channel from one rib of a rib-pair on the other rib of the ribs Pair is headed.
  • a greatly increased turbulence of the cooling air a greatly increased local heat transfer coefficient is connected, so that overall, compared to known solutions, a significantly more effective cooling, in particular in the region of the flow inlet edge, can be provided.
  • the turbine blade according to the invention can thus be exposed to higher gas temperatures, even if no film cooling is provided. If film cooling is provided, even higher gas temperatures are possible.
  • a high degree of turbulence is formed on the flowed-on ribs, which, in combination with impingement cooling effects and a strong increase in surface area on the cooling air side, leads to efficient use of cooling air and equalization of the temperature distribution.
  • the two ribs of a rib pair are arranged in the cooling channel such that a flow of a cooling medium flowing in the cooling channel, preferably in the form of cooling air, from one rib of the rib pair essentially transversely to the other rib of the rib pair.
  • the two ribs of a rib pair include a predetermined angle, and a total cooling capacity of the two ribs of a rib pair is adjusted over the angle to a predetermined cooling requirement for the flow inlet edge in the vicinity of the rib pair.
  • the extent of the turbulence of the cooling air and thus also the local heat transfer coefficient can be selectively influenced, so that cooling adapted to a local cooling requirement for the flow inlet edge can be realized.
  • the cooling capacity of a pair of ribs can be increased by increasing the angle enclosed by the two ribs of the ribbed pair.
  • the temperature distribution at the flow inlet edge can be "made uniform" by means of this practical development, since according to the invention comparatively hot spots of the flow inlet edge by appropriately trained rib pairs a correspondingly strong cooling and vice versa, so that an effective cooling of the flow inlet edge can be realized which counteracts an inhomogeneous temperature distribution.
  • An inhomogeneous temperature distribution is associated with high thermal loads, which adversely affect the life of the turbine. This applies in particular to turbine blades which are used in axially through-flow turbines in which an inhomogeneous temperature distribution along the radial direction is formed for the flow inlet edge.
  • the ribs extend projecting from a wall bounding the cooling channel into the cooling channel, the ribs preferably being formed integrally with the bounding wall.
  • the rib pairs are mounted within an insert which is inserted into the cooling channel.
  • an insert is provided according to the invention, which can optionally be removed from the turbine blade, preferably in the form of a guide vane, to adapt, for example, the angular position of the rib pairs of a given application.
  • the casting of the turbine blade can also be kept simple, so that the turbine blade according to the invention can also be produced without elaborately designed casting cores.
  • the cooling channel extends parallel to the flow inlet edge continuously through the turbine blade to provide effective cooling along the entire extent of the flow inlet edge.
  • FIG. 1 shows a sketch-like sectional view of a turbine blade 10 according to the invention through its flow inlet edge 12.
  • the section according to the sectional surface A-A of FIG. 1 is shown in FIG. 3, this being a sketch-like sectional view of the front section of a turbine blade 10 according to the invention.
  • a cooling channel 14 extending parallel to the flow inlet edge 12 is formed near the flow inlet edge 12 (ie, a radially extending channel 14 in the case of axially through-flowed turbines).
  • a number of pairs of ribs 24 are arranged in succession in this, the individual ribs 18 of each rib pair 24 being set transversely to each other by a predetermined angle ⁇ .
  • the ribs 18 of a pair of ribs 24, viewed along the cooling channel extension, are arranged offset to one another.
  • the ribs 18 of each pair 24 and the ribs 18 of immediately adjacent pairs 24 are thus arranged overlapping in skating step shape.
  • the cooling air When flowing through the cooling channel 14, the cooling air is alternately deflected by the individual ribs 18 of each pair 24. A high degree of turbulence is formed on the flowed-on ribs 18, which, in combination with impingement cooling effects and the associated cooling air-side surface enlargement, leads to an efficient use of cooling air.
  • the angle is ⁇ in larger than in the edge regions of the turbine blade 10, so as to cool the middle, during operation usually strongly heated area of the flow inlet edge 12 stronger than the edge regions of the flow inlet edge 12.
  • Suitable values for the angle ⁇ which are adapted to the respective cooling requirement, according to the invention are in the range of about 60 ° to 90 °.
  • the individual ribs 18 of a pair 24 extend predominantly from a front wall Alternatively, however, the ribs 18 may be fixed on one side only to the front wall 16 without extending to the rear wall 20. Likewise, the ribs may also be part of an insert which can be inserted in the cooling channel 14.
  • the cooling air can preferably be guided in the direction of the front wall 16 by suitably setting the angular position ⁇ , in order to achieve the most effective possible cooling of the flow inlet edge 12.
  • provided angular sizes are in the range of about 30 ° to 60 °.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Architecture (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP06023377A 2006-11-09 2006-11-09 Turbinenschaufel Withdrawn EP1921269A1 (de)

Priority Applications (5)

Application Number Priority Date Filing Date Title
EP06023377A EP1921269A1 (de) 2006-11-09 2006-11-09 Turbinenschaufel
US12/513,682 US8215909B2 (en) 2006-11-09 2007-10-18 Turbine blade
JP2009535661A JP5329418B2 (ja) 2006-11-09 2007-10-18 タービン翼
PCT/EP2007/061127 WO2008055764A1 (de) 2006-11-09 2007-10-18 Turbinenschaufel
EP07821492.1A EP2087207B1 (de) 2006-11-09 2007-10-18 Turbinenschaufel

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP06023377A EP1921269A1 (de) 2006-11-09 2006-11-09 Turbinenschaufel

Publications (1)

Publication Number Publication Date
EP1921269A1 true EP1921269A1 (de) 2008-05-14

Family

ID=37909821

Family Applications (2)

Application Number Title Priority Date Filing Date
EP06023377A Withdrawn EP1921269A1 (de) 2006-11-09 2006-11-09 Turbinenschaufel
EP07821492.1A Not-in-force EP2087207B1 (de) 2006-11-09 2007-10-18 Turbinenschaufel

Family Applications After (1)

Application Number Title Priority Date Filing Date
EP07821492.1A Not-in-force EP2087207B1 (de) 2006-11-09 2007-10-18 Turbinenschaufel

Country Status (4)

Country Link
US (1) US8215909B2 (enrdf_load_stackoverflow)
EP (2) EP1921269A1 (enrdf_load_stackoverflow)
JP (1) JP5329418B2 (enrdf_load_stackoverflow)
WO (1) WO2008055764A1 (enrdf_load_stackoverflow)

Families Citing this family (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20120076660A1 (en) * 2010-09-28 2012-03-29 Spangler Brandon W Conduction pedestals for a gas turbine engine airfoil
US8920122B2 (en) 2012-03-12 2014-12-30 Siemens Energy, Inc. Turbine airfoil with an internal cooling system having vortex forming turbulators
WO2018153796A1 (en) * 2017-02-24 2018-08-30 Siemens Aktiengesellschaft A turbomachine blade or vane having a cooling channel with a criss-cross arrangement of pins
GB2574368A (en) * 2018-04-09 2019-12-11 Rolls Royce Plc Coolant channel with interlaced ribs
US10787932B2 (en) 2018-07-13 2020-09-29 Honeywell International Inc. Turbine blade with dust tolerant cooling system
US10989067B2 (en) 2018-07-13 2021-04-27 Honeywell International Inc. Turbine vane with dust tolerant cooling system
US10669862B2 (en) * 2018-07-13 2020-06-02 Honeywell International Inc. Airfoil with leading edge convective cooling system
GB201902997D0 (en) 2019-03-06 2019-04-17 Rolls Royce Plc Coolant channel
US11230929B2 (en) 2019-11-05 2022-01-25 Honeywell International Inc. Turbine component with dust tolerant cooling system

Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5395212A (en) * 1991-07-04 1995-03-07 Hitachi, Ltd. Member having internal cooling passage
US5472316A (en) * 1994-09-19 1995-12-05 General Electric Company Enhanced cooling apparatus for gas turbine engine airfoils
DE19526917A1 (de) * 1995-07-22 1997-01-23 Fiebig Martin Prof Dr Ing Längswirbelerzeugende Rauhigkeitselemente
US5695321A (en) * 1991-12-17 1997-12-09 General Electric Company Turbine blade having variable configuration turbulators
EP1380724A2 (en) * 2002-07-11 2004-01-14 Mitsubishi Heavy Industries, Ltd. Cooled turbine blade
EP1637699A2 (en) * 2004-09-09 2006-03-22 General Electric Company Offset coriolis turbulator blade

Family Cites Families (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4416585A (en) * 1980-01-17 1983-11-22 Pratt & Whitney Aircraft Of Canada Limited Blade cooling for gas turbine engine
JP3396360B2 (ja) * 1996-01-12 2003-04-14 三菱重工業株式会社 ガスタービン冷却動翼
DE19634238A1 (de) * 1996-08-23 1998-02-26 Asea Brown Boveri Kühlbare Schaufel
US5797726A (en) * 1997-01-03 1998-08-25 General Electric Company Turbulator configuration for cooling passages or rotor blade in a gas turbine engine
EP0892149B1 (de) * 1997-07-14 2003-01-22 ALSTOM (Switzerland) Ltd Kühlsystem für den Vorderkantenbereich einer hohlen Gasturbinenschaufel
EP1191189A1 (de) * 2000-09-26 2002-03-27 Siemens Aktiengesellschaft Gasturbinenschaufel
US8690538B2 (en) * 2006-06-22 2014-04-08 United Technologies Corporation Leading edge cooling using chevron trip strips
US20070297916A1 (en) * 2006-06-22 2007-12-27 United Technologies Corporation Leading edge cooling using wrapped staggered-chevron trip strips

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5395212A (en) * 1991-07-04 1995-03-07 Hitachi, Ltd. Member having internal cooling passage
US5695321A (en) * 1991-12-17 1997-12-09 General Electric Company Turbine blade having variable configuration turbulators
US5472316A (en) * 1994-09-19 1995-12-05 General Electric Company Enhanced cooling apparatus for gas turbine engine airfoils
DE19526917A1 (de) * 1995-07-22 1997-01-23 Fiebig Martin Prof Dr Ing Längswirbelerzeugende Rauhigkeitselemente
EP1380724A2 (en) * 2002-07-11 2004-01-14 Mitsubishi Heavy Industries, Ltd. Cooled turbine blade
EP1637699A2 (en) * 2004-09-09 2006-03-22 General Electric Company Offset coriolis turbulator blade

Also Published As

Publication number Publication date
EP2087207A1 (de) 2009-08-12
EP2087207B1 (de) 2016-04-20
US20100054952A1 (en) 2010-03-04
WO2008055764A1 (de) 2008-05-15
US8215909B2 (en) 2012-07-10
JP5329418B2 (ja) 2013-10-30
JP2010509535A (ja) 2010-03-25

Similar Documents

Publication Publication Date Title
EP2087206B1 (de) Turbinenschaufel
EP2087207B1 (de) Turbinenschaufel
EP1223308B1 (de) Komponente einer Strömungsmaschine
EP1113145B1 (de) Schaufel für Gasturbinen mit Drosselquerschnitt an Hinterkante
DE60018817T2 (de) Gekühlte Gasturbinenschaufel
DE60025074T2 (de) Methode zur Kühlung einer Wand einer Strömungsmaschinenschaufel
EP2828484B2 (de) Turbinenschaufel
EP1320661B1 (de) Gasturbinenschaufel
DE3345263C2 (de) Gekühlte Turbinenschaufel
DE602005000350T2 (de) Turbinenstatorschaufel mit verbesserter Kühlung
EP2611990B1 (de) Turbinenschaufel für eine gasturbine
EP3658751B1 (de) Schaufelblatt für eine turbinenschaufel
DE102011000878B4 (de) Turbinenschaufel mit abgeschirmtem Kühlmittelzuführungskanal
EP0798448A2 (de) Vorrichtung und Verfahren zur Kühlung einer einseitig von Heissgas umgebenen Wand
DE2060123A1 (de) Turbinenrotorschaufel mit vom Schaufelfuss zum Schaufelende verlaufenden Kuehlkanaelen
WO2008092725A1 (de) Turbinenschaufel
CH705185A1 (de) Schaufel für eine Gasturbine sowie Verfahren zum Herstellen einer solchen Schaufel.
EP1207269B1 (de) Gasturbinenschaufel
EP3087254B1 (de) Heissgasbeaufschlagbares bauteil für eine gasturbine sowie dichtungsanordnung mit einem derartigen bauteil
EP3232001A1 (de) Laufschaufel für eine turbine
DE112015003273T5 (de) Abgasturbolader
EP1445423B1 (de) Kühlbare Turbomaschinenschaufel
EP0992653A1 (de) Filmgekühlte Komponenten mit Filmkühlungkanälen mit dreieckigem Querschnitt
DE19845147B4 (de) Vorrichtung und Verfahren zur Kühlung einer einseitug von Heißgas umgebenen Wand
EP1998115A1 (de) Kühlkanal zum Kühlen einer ein Heißgas führenden Komponente

Legal Events

Date Code Title Description
PUAI Public reference made under article 153(3) epc to a published international application that has entered the european phase

Free format text: ORIGINAL CODE: 0009012

AK Designated contracting states

Kind code of ref document: A1

Designated state(s): AT BE BG CH CY CZ DE DK EE ES FI FR GB GR HU IE IS IT LI LT LU LV MC NL PL PT RO SE SI SK TR

AX Request for extension of the european patent

Extension state: AL BA HR MK RS

AKX Designation fees paid
STAA Information on the status of an ep patent application or granted ep patent

Free format text: STATUS: THE APPLICATION IS DEEMED TO BE WITHDRAWN

18D Application deemed to be withdrawn

Effective date: 20081115

REG Reference to a national code

Ref country code: DE

Ref legal event code: 8566